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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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At least 145 records · Page 8
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NASA NTRS2013

Aqua 10 Years After Launch

A little over ten years ago, in the early morning hours of May 4, 2002, crowds of spectators stood anxiously watching as the Delta II rocket carrying NASA's Aqua spacecraft lifted off from its launch pad at Vandenberg Air Force Base in California at 2:55 a.m. The rocket quickly went through a low-lying cloud cover, after which the main portion of the rocket fell to the waters below and the rockets second stage proceeded to carry Aqua south across the Pacific, onward over Antarctica, and north to Africa, where the spacecraft separated from the rocket 59.5 minutes after launch. Then, 12.5 minutes later, the solar array unfurled over Europe, and Aqua was on its way in the first of what by now have become over 50,000 successful orbits of the Earth.

Parkinson, Claire L.↗
NASA NTRS2012

Simulation Study of a Follow-on Gravity Mission to GRACE

The gravity recovery and climate experiment (GRACE) has been providing monthly estimates of the Earth's time-variable gravity field since its launch in March 2002. The GRACE gravity estimates are used to study temporal mass variations on global and regional scales, which are largely caused by a redistribution of water mass in the Earth system. The accuracy of the GRACE gravity fields are primarily limited by the satellite-to-satellite range-rate measurement noise, accelerometer errors, attitude errors, orbit errors, and temporal aliasing caused by unmodeled high-frequency variations in the gravity signal. Recent work by Ball Aerospace and Technologies Corp., Boulder, CO has resulted in the successful development of an interferometric laser ranging system to specifically address the limitations of the K-band microwave ranging system that provides the satellite-to-satellite measurements for the GRACE mission. Full numerical simulations are performed for several possible configurations of a GRACE Follow-On (GFO) mission to determine if a future satellite gravity recovery mission equipped with a laser ranging system will provide better estimates of time-variable gravity, thus benefiting many areas of Earth systems research. The laser ranging system improves the range-rate measurement precision to approximately 0.6 nm/s as compared to approx. 0.2 micro-seconds for the GRACE K-band microwave ranging instrument. Four different mission scenarios are simulated to investigate the effect of the better instrument at two different altitudes. The first pair of simulated missions is flown at GRACE altitude (approx. 480 km) assuming on-board accelerometers with the same noise characteristics as those currently used for GRACE. The second pair of missions is flown at an altitude of approx. 250 km which requires a drag-free system to prevent satellite re-entry. In addition to allowing a lower satellite altitude, the drag-free system also reduces the errors associated with the accelerometer. All simulated mission scenarios assume a two satellite co-orbiting pair similar to GRACE in a near-polar, near-circular orbit. A method for local time variable gravity recovery through mass concentration blocks (mascons) is used to form simulated gravity estimates for Greenland and the Amazon region for three GFO configurations and GRACE. Simulation results show that the increased precision of the laser does not improve gravity estimation when flown with on-board accelerometers at the same altitude and spacecraft separation as GRACE, even when time-varying background models are not included. This study also shows that only modest improvement is realized for the best-case scenario (laser, low-altitude, drag-free) as compared to GRACE due to temporal aliasing errors. These errors are caused by high-frequency variations in the hydrology signal and imperfections in the atmospheric, oceanographic, and tidal models which are used to remove unwanted signal. This work concludes that applying the updated technologies alone will not immediately advance the accuracy of the gravity estimates. If the scientific objectives of a GFO mission require more accurate gravity estimates, then future work should focus on improvements in the geophysical models, and ways in which the mission design or data processing could reduce the effects of temporal aliasing.

Loomis, Bryant D.↗
NASA NTRS2012

Note: Silicon Carbide Telescope Dimensional Stability for Space-based Gravitational Wave Detectors

Space-based gravitational wave detectors are conceived to detect gravitational waves in the low frequency range by measuring the distance between proof masses in spacecraft separated by millions of kilometers. One of the key elements is the telescope which has to have a dimensional stability better than 1 pm Hz(exp −1/2) at 3 mHz. In addition, the telescope structure must be light, strong, and stiff. For this reason a potential telescope structure consisting of a silicon carbide quadpod has been designed, constructed, and tested. We present dimensional stability results meeting the requirements at room temperature. Results at −60 C are also shown although the requirements are not met due to temperature fluctuations in the setup.

LISA-like↗
NASA NTRS2014

Three Orbital Burns to Molniya Orbit via Shuttle Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982-86 joint NASA-USAF Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57deg inclined orbit: 9,545 lb vs. 9,552 lb of separated spacecraft weight respectively. There was a significant reduction in the need for propellant launch time reserve for a one hour window: only 78 lb for the three burn transfer vs. 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three vs. two burn transfer (12 vs. 11/4 hrs), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗
NASA NTRS2015

Three Orbital Burns to Molniya Orbit Via Shuttle_Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982 to 1986 joint National Aeronautics and Space Administration (NASA)-United States Air Force (USAF) Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37 deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57 deg inclined orbit: 9,545 versus 9,552 lb of separated spacecraft weight, respectively. There was a significant reduction in the need for propellant launch time reserve for a 1 hr window: only 78 lb for the three burn transfer versus 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three versus two burn transfer (12 vs. 1-1/4 hr), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗
NASA NTRS2013

Optimization of the Orbiting Wide-Angle Light Collectors (OWL) Mission for Charged-Particle and Neutrino Astronomy

OWL [1] uses the Earth's atmosphere as a vast calorimeter to fully enable the emerging field of charged-particle astronomy with high-statistics measurements of ultra-high-energy cosmic rays (UHECR) and a search for sources of UHE neutrinos and photons. Confirmation of the Greisen-Zatsepin-Kuzmin (GZK) suppression above approx. 4 x 10(exp 19) eV suggests that most UHECR originate in astrophysical objects. Higher energy particles must come from sources within about 100 Mpc and are deflected by approx. 1 degree by predicted intergalactic/galactic magnetic fields. The Pierre Auger Array, Telescope Array and the future JEM-EUSO ISS mission will open charged-particle astronomy, but much greater exposure will be required to fully identify and measure the spectra of individual sources. OWL uses two large telescopes with 3 m optical apertures and 45 degree FOV in near-equatorial orbits. Simulations of a five-year OWL mission indicate approx. 10(exp 6) sq km/ sr/ yr of exposure with full aperture at approx. 6 x 10(exp 19) eV. Observations at different altitudes and spacecraft separations optimize sensitivity to UHECRs and neutrinos. OWL's stereo event reconstruction is nearly independent of track inclination and very tolerant of atmospheric conditions. An optional monocular mode gives increased reliability and can increase the instantaneous aperture. OWL can fully reconstruct horizontal and upward-moving showers and so has high sensitivity to UHE neutrinos. New capabilities in inflatable structures optics and silicon photomultipliers can greatly increase photon sensitivity, reducing the energy threshold for n detection or increasing viewed area using a higher orbit. Design trades between the original and optimized OWL missions and the enhanced science capabilities are described.

Optimization↗
NASA NTRS2015

FJET Database Project: Extract, Transform, and Load

The Data Mining & Knowledge Management team at Kennedy Space Center is providing data management services to the Frangible Joint Empirical Test (FJET) project at Langley Research Center (LARC). FJET is a project under the NASA Engineering and Safety Center (NESC). The purpose of FJET is to conduct an assessment of mild detonating fuse (MDF) frangible joints (FJs) for human spacecraft separation tasks in support of the NASA Commercial Crew Program. The Data Mining & Knowledge Management team has been tasked with creating and managing a database for the efficient storage and retrieval of FJET test data. This paper details the Extract, Transform, and Load (ETL) process as it is related to gathering FJET test data into a Microsoft SQL relational database, and making that data available to the data users. Lessons learned, procedures implemented, and programming code samples are discussed to help detail the learning experienced as the Data Mining & Knowledge Management team adapted to changing requirements and new technology while maintaining flexibility of design in various aspects of the data management project.

excel vba↗
NASA NTRS2015

Towing Asteroids with Gravity Tractors Enhanced by Tethers and Solar Sails

Material collected from an asteroid's surface can be used to increase gravitational attraction between the asteroid and a Gravity Tractor (GT); the spacecraft therefore operates more effectively and is referred to as an Enhanced Gravity Tractor (EGT). The use of tethers and solar sails to further improve effectiveness and simplify operations is investigated. By employing a tether, the asteroidal material can be placed close to the asteroid while the spacecraft is stationed farther away, resulting in a better safety margin and improved thruster efficiency. A solar sail on a spacecraft can naturally provide radial offset and inter-spacecraft separation required for multiple EGTs.

Shen, Haijun↗
NASA NTRS2017

Comparative Study of Three Reconnection X Line Models at the Earth's Dayside Magnetopause Using in Situ Observations

This work examines the large-scale aspects of magnetic field reconnection at the Earth's dayside magnetopause. We use two sets of reconnection events, which are identified mostly by the in situ detection of accelerated and Alfvenic plasma flows. We intercompare three analytical models that predict the reconnection X line location and orientation, namely, the Trattner et al. (2007) and Swisdak and Drake (2007) models and also a modified version of the component merging model. In the first set of reconnection observations, we show three fortuitous, quasi-simultaneous dayside magnetopause crossing events where two widely separated spacecraft detect reconnection signatures, and the X line location and orientation can be inferred from the observations. We compare X line model predictions to those inferred from observations. These three reconnection events indicate the presence of an extended (greater than 7 Earth radii in length), component-type reconnection X line on Earth's dayside magnetopause connecting and structuring the reconnection signatures at locations far apart. In the second set of reconnection events, we analyze the X line models' performance in predicting the observed reconnection outflow direction, i.e., its north-south and/or east-west senses, in a total of 75 single, rather than multiple and quasi-simultaneous, magnetopause crossing events, where reconnection-associated plasma flows were clearly present. We found that the Swisdak and Drake's (2007) X line model performs slightly better, albeit not statistically significant, when predicting both accelerated plasma flow north-south and east-west components in 73% and 53% of the cases, respectively, as compared to the Trattner et al. (2007) model (70% north-south and 42% east-west) and the modified component merging model (66% north-south and 50% east-west).

Souza, V. M.↗
NASA NTRS2018

A Small Spacecraft Swarm Deployment and Stationkeeping Strategy for Sun-Earth L1 Halo Orbits

Spacecraft orbits about the Sun-Earth librarian point L1 have been of interest since the 1950s. An L1 halo orbit was first achieved with the International Sun-Earth Explorer-3 (ISEE-3) mission, and similar orbits around Sun-Earth L1 were achieved in the Solar and Heliospheric Observatory (SOHO), Advanced Composition Explorer (ACE), Genesis, and Deep Space Climate Observatory (DSCOVR) missions. With recent advancements in CubeSat technology, we envision that it will soon be feasible to deploy CubeSats at L1. As opposed to these prior missions where one large satellite orbited alone, a swarm of CubeSats at L1 would enable novel science data return, providing a topology for intersatellite measurements of heliophysics phenomena both spatially and temporally, at varying spatial scales.The purpose of this iPoster is to present a flight dynamics strategy for a swarm of numerous CubeSats orbiting Sun-Earth L1. The presented method is a coupled, two-part solution. First, we present a deployment strategy for theCubeSats that is optimized to produce prescribed, time-varying intersatellite baselines for the purposes of collectingmagnetometer data as well as radiometric measurements from cross-links. Second, we employ a loose controlstrategy that was successfully applied to SOHO and ACE for minimized stationkeeping propellant expenditure. Weemphasize that the presented solution is practical within the current state-of-the-art and heritage CubeSat technology,citing capabilities of CubeSat designs that will launch on the upcoming Exploration Mission 1 (EM-1) to lunar orbitsand beyond. Within this iPoster, we present animations of the simulated deployment strategy and resulting spacecrafttrajectories. Mission design parameters such as total v required for long-term station keeping andminimum/maximum/mean spacecraft separation distances are also presented.

Optimal control↗
NASA NTRSDate not supplied

TPSAS-NF1676L-14919-DND

Manned spacecraft abort during ascent is a particularly unpleasant event. Not only does this represent a mission failure, but it can also be a very severe loading event on the crew. There are several dynamic events happening within a few hundred milliseconds while the vehicle is under aerodynamic-induced strain and possibly divergent control. Some of these events are spacecraft separation, abort motor ignitions, and aerodynamic variation. This paper will describe a methodology for implementing this analysis within a multi-body solution. Some generic examples will be shown along with some guidance on what parameters are important.

Isam S Yunis↗
NASA NTRS2019

A Realistic Roadmap to Formation Flying Space Interferometry

The ultimate astronomical observatory would be a formation flying interferometer in space, immune to atmospheric turbulence and absorption, free from atmospheric and telescope thermal emission, and reconfigurable to adjust baselines according to the required angular resolution. Imagine the near/mid-infrared sensitivity of the JWST and the far-IR sensitivity of Herschel but with ALMA-level angular resolution, or imagine having the precision control to null host star light across 250m baselines and to detect molecules from the atmospheres of nearby exo-Earths. With no practical engineering limit to the formation’s size or number of telescopes in the array, formation flying interferometry will revolutionize astronomy and this White Paper makes the case that it is now time to accelerate investments in this technological area. Here we provide a brief overview of the required technologies needed to allow light to be collected and interfered using separate spacecrafts. We emphasize the emerging role of inexpensive smallSat projects and the excitement for the LISA Gravitational Wave Interferometer to push development of the required engineering building-blocks. We urge the Astro2020 Decadal Survey Committee to highlight the need for a small-scale formation flying space interferometer project to demonstrate end-to-end competency with a timeline for first stellar fringes by the end of the decade.

Gioia Rau↗
NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of these samples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system. The results of these studies will also be examined in this paper.

Darius Yaghoubi↗
NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of thesesamples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system.

Mars↗
NASA NTRS2002

Simple approach to laser frequency stabilization

We describe a simple approach to laser frequency stabilization. It was originally developed for the metrology subsystem for NASA's StarLight mission, a space-based separated-spacecraft stellar interferometer.

starlight↗
NASA NTRS1995

Technology for Space Optical Interferometry

Optical interferometers are expected to provide the next great leap forward in space based astronomy beyond the Hubble Space Telescope. Various systems are envisioned, ranging from large orbiting structures to lunar based facilities to separated spacecraft flying in formation. Regardless of the specific architecture, these systems will present unprecedented challenges in the measurement and control of optical surfaces, hence driving the technological state-of-the-art in the areas of laser metrology and alignment and stabilization of optomechanical systems.

interferometry↗
NASA NTRS1998

The LISA Mission Design

The Laser Interferometer Space Antenna (LISA) will be capable of detecting gravitational waves with frequencies from 0.1 mHz to 1 Hz by using laser interferometers to monitor changes in the distances between test masses in spacecraft separated by five million km.

Laser↗