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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Systems Integration Processes for NASA Ares I Crew Launch Vehicle

NASA's Exploration Initiative will require development of many new elements to constitute a robust system of systems. New launch vehicles are needed to place cargo and crew in stable Low Earth Orbit (LEO). This paper examines the systems integration processes NASA is utilizing to ensure integration and control of propulsion and nonpropulsion elements within NASA's Crew Launch Vehicle (CLV), now known as the Ares I. The objective of the Ares I is to provide the transportation capabilities to meet the Constellation Program requirements for delivering a Crew Exploration Vehicle (CEV) or other payload to LEO in support of the lunar and Mars missions. The Ares I must successfully provide this capability within cost and schedule, and with an acceptable risk approach. This paper will describe the systems engineering management processes that will be applied to assure Ares I Project success through complete and efficient technical integration. Discussion of technical review and management processes for requirements development and verification, integrated design and analysis, integrated simulation and testing, and the integration of reliability, maintainability and supportability (RMS) into the design will also be included. The Ares I Project is logically divided into elements by the major hardware groupings, and associated management, system engineering, and integration functions. The processes to be described herein are designed to integrate within these Ares I elements and among the other Constellation projects. Also discussed is launch vehicle stack integration (Ares I to CEV, and Ground and Flight Operations integration) throughout the life cycle, including integrated vehicle performance through orbital insertion, recovery of the first stage, and reentry of the upper stage. The processes for decomposing requirements to the elements and ensuring that requirements have been correctly validated, decomposed, and allocated, and that the verification requirements are properly defined to ensure that the system design meets requirements, will be discussed.

Taylor, James L.↗
NASA NTRS2015

A Future Large-Aperture UVOIR Space Observatory: Reference Designs

Our joint NASA GSFC/JPL/MSFC/STScI study team has used community-provided science goals to derive mission needs, requirements, and candidate mission architectures for a future large-aperture, non-cryogenic UVOIR space observatory. We describe the feasibility assessment of system thermal and dynamic stability for supporting coronagraphy. The observatory is in a Sun-Earth L2 orbit providing a stable thermal environment and excellent field of regard. Reference designs include a 36-segment 9.2 m aperture telescope that stows within a five meter diameter launch vehicle fairing. Performance needs developed under the study are traceable to a variety of reference designs including options for a monolithic primary mirror.

exo-planets↗
NASA NTRSDate not supplied

Development of a CubeSat-Scale Green Monopropellant Propulsion System for NASA’s Lunar Flashlight Mission

NASA’s Lunar Flashlight is a low-cost 6U CubeSat whose mission is to search for ice and mineral deposits inside of the scattered craters at Moon’s southern pole. To conduct its primary science mission, Lunar Flashlight must be placed in a stable lunar polar orbit which requires the utilization of an on-board propulsion system. However, to this date, most CubeSats have been propelled by cold-gas or electric propulsion systems that have proven to scale well but lack sufficient impulse to conduct large ΔV maneuvers such as orbit insertions. To this end, the Lunar Flashlight mission has chosen to utilize a custom-designed green monopropellant propulsion system developed by the Georgia Institute of Technology under the leadership ofNASA’s Marshall Space Flight Center and support from the Jet Propulsion Laboratory. The developed system is capable of providing more than the required propulsive capability for full mission success while fitting inside of a 2.5U volume and weighing less than six kilograms. The system utilizes the Advanced Spacecraft Energetic Non Toxic (ASCENT) green monopropellant that provides higher specific impulse compared to traditional hydrazine while also being safer to handle. If successful, the presented propulsion system will enable Lunar Flashlight to be the first CubeSat to reach the Moon, the first to conduct an orbit insertion, and will be the firstCubeSat demonstration of the ASCENT propellant.

Grayson Huggins↗
NASA NTRS2021

Development of a CubeSat-Scale Green Monopropellant Propulsion System for NASA's Lunar Flashlight Mission

NASA's Lunar Flashlight is a low-cost 6U CubeSat whose mission is to search for ice and mineral deposits inside of the scattered craters at Moon's southern pole. To conduct its primary science mission, Lunar Flashlight must be placed in a stable lunar polar orbit which requires the utilization of an on-board propulsion system. However, to this date, most CubeSats have been propelled by cold-gas or electric propulsion systems that have proven to scale well but lack sufficient impulse to conduct large $\Delta V$ maneuvers such as orbit insertions. To this end, the Lunar Flashlight mission has chosen to utilize a custom-designed green monopropellant propulsion system developed by the Georgia Institute of Technology under the leadership of NASA's Marshall Space Flight Center and support from the Jet Propulsion Laboratory. The developed system is capable of providing more than the required propulsive capability for full mission success while fitting inside of a 2.5U volume and weighing less than six kilograms. The system utilizes the Advanced Space Craft Energetic Non Toxic (ASCENT) green monopropellant that provides higher specific impulse compared to traditional hydrazine while also being safer to handle. If successful, the presented propulsion system will enable Lunar Flashlight to be the first CubeSat to reach the Moon, the first to conduct an orbit insertion, and will be the first CubeSat demonstration of the ASCENT propellant.

Kowalkowski, Matthew↗
NASA NTRS1982

Periodic orbits of the elliptic restricted problem for the Sun-Jupiter-Saturn system

A systematic approach to generate periodic orbits in the elliptic restricted problem of three bodies is introduced. The approach is based on (numerical) continuation from periodic orbits of the first and second kind in the circular restricted problem to periodic orbits in the elliptic restricted problem. Two families of periodic orbits of the elliptic restricted problem are found by this approach. The mass ratio of the primaries of these orbits is equal to that of the Sun-Jupiter system. The sidereal mean motions between the infinitesimal body and the smaller primary are in a 2:5 resonance, so as to approximate the Sun-Jupiter-Saturn system. The lineaar stabilities of these periodic orbits are studied as functions of the eccentricities of the primaries and of the infinitesimal body. The results show that both stable and unstable periodic orbits exist in the elliptic restricted problem that are close to the actual Sun-Jupiter-Saturn system. However, the periodic orbit closest to the actual Sun-Jupiter-Saturn system is (linearly) stable.

Kwok, J. H.↗
NASA NTRS1995

Orbital and Landing Operations at Near-Earth

Orbital and landing operations about near-Earth asteroids are different than classical orbital operations about large bodies. The major differences lie with the small mass of the asteroid, the lower orbital velocities, the larger Solar tide and radiation pressure perturbations, the irregular shape of the asteroid and the potential for non-uniform rotation of the asteroid. These differences change the nature of orbits about an asteroid to where it is often common to find trajectories that evolve from stable, near-circular orbits to crashing or escaping orbits in a matter of days. The understanding and control of such orbits is important if a human or robotic presence at asteroids is to be commonplace in the future.

Orbital Near-Earth↗
NASA NTRS1991

Pinhole cameras as sensors for atomic oxygen in orbit; application to attitude determination of the LDEF

Images produced by pinhole cameras using film sensitive to atomic oxygen provide information on the ratio of spacecraft orbital velocity to the most probable thermal speed of oxygen atoms, provided the spacecraft orientation is maintained stable relative to the orbital direction. Alternatively, as it is described, information on the spacecraft attitude relative to the orbital velocity can be obtained, provided that corrections are properly made for thermal spreading and a co-rotating atmosphere. The LDEF orientation, uncorrected for a co-rotating atmosphere, was determined to be yawed 8.0 minus/plus 0.4 deg from its nominal attitude, with an estimated minus/plus 0.35 deg oscillation in yaw. The integrated effect of inclined orbit and co-rotating atmosphere produces an apparent oscillation in the observed yaw direction, suggesting that the LDEF attitude measurement will indicate even better stability when corrected for a co-rotating atmosphere. The measured thermal spreading is consistent with major exposure occurring during high solar activity, which occurred late during the LDEF mission.

Peters, Palmer N.↗
NASA NTRS1992

Pinhole cameras as sensors for atomic oxygen in orbit: Application to attitude determination of the LDEF

Images produced by pinhole cameras using film sensitive to atomic oxygen provide information on the ratio of spacecraft orbital velocity to the most probable thermal speed of oxygen atoms, provided the spacecraft orientation is maintained stable relative to the orbital direction. Alternatively, information on the spacecraft attitude relative to the orbital velocity can be obtained, provided that corrections are properly made for thermal spreading and a corotating atmosphere. The Long Duration Exposure Facility (LDEF) orientation, uncorrected for a corotating atmosphere, was determined to be yawed 8.0 +/- 0.4 degrees from its nominal attitude, with an estimated +/- 0.35 degree oscillation in yaw. The integrated effect of inclined orbit and corotating atmosphere produces an apparent oscillation in the observed yaw direction, suggesting that the LDEF attitude measurement will indicate even better stability when corrected for a corotating atmosphere. The measured thermal spreading is consistent with major exposure occurring during high solar activity, which occurred late during the LDEF mission.

Peters, Palmer N.↗
NASA NTRS2012

The Precise Orbit and the Challenge of Long Term Stability

The computation of a precise orbit reference is a fundamental component of the altimetric measurement. Since the dawn of the modern altimeter age, orbit accuracy has been determined by the quality of the GPS, SLR, and DORIS tracking systems, the fidelity of the measurement and force models, and the choice of parameterization for the orbit solutions, and whether a dynamic or a reduced-dynamic strategy is used to calculate the orbits. At the start of the TOPEX mission, the inaccuracies in the modeling of static gravity, dynamic ocean tides, and the nonconservative forces dominated the orbit error budget. Much of the error due to dynamic mismodeling can be compensated by reduced-dynamic tracking techniques depending on the measurement system strength. In the last decade, the launch of the GRACE mission has eliminated the static gravity field as a concern, and the background force models and the terrestrial reference frame have been systematically refined. GPS systems have realized many improvements, including better modeling of the forces on the GPS spacecraft, large increases in the ground tracking network, and improved modeling of the GPS measurements. DORIS systems have achieved improvements through the use of new antennae, more stable monumentation, and of satellite receivers that can track multiple beacons, and as well as through improved modeling of the nonconservative forces. Many of these improvements have been applied in the new reprocessed time series of orbits produced for the ERS satellites, Envisat, TOPEX/Poseidon and the Jason satellites, and as well as for the most recent Cryosat-2 and HY2A. We now face the challenge of maintaining a stable orbit reference for these altimetric satellites. Changes in the time-variable gravity field of the Earth and how these are modelled have been shown to affect the orbit evolution, and the calibration of the altimetric data with tide gauges. The accuracy of the reference frame realizations, and their projection into the future remains a source of error. Other sources of omission error include the geocenter for which no consensus model is as of yet applied. Although progress has been made in nonconservative force modeling through the use of detailed satellite-specific models, radiation pressure modeling, and atmospheric density modeling remain a potential source of orbit error. The longer term influence of variations in the solar and terrestrial radiation fields over annual and solar cycles remains principally untested. Also the long term variation in optical and thermal properties of the space vehicle surfaces would contribute to biases in the orbital frame if ignored. We review the status of altimetric precision orbit determination as exemplified by the recent computations undertaken by the different analysis centers for ERS, Envisat, TOPEX/Poseidon, Jason, Cryosat2 and HY2A, and we provide a perspective on the challenges for future missions such as the Jason-3, SENTINEL-3 and SWOT.

Lemoine, Frank G.↗
NASA NTRS1974

Periodic Trojan-type orbits in the earth-sun system

Periodic orbits about the triangular equilibrium points are found for the planar restricted three-body problem using the earth-sun system. The maximum semimajor axis for tadpole orbits ranges from the infinitesimal orbit at 1.000 AU to the near-limiting orbit at 1.00285 AU. Horseshoe orbits are found for 1.0029 to 1.0080 AU, larger horseshoes being unstable because of close approaches to the earth. Using stability tests devised by Rabe (1961, 1962), the limit of stability for nonperiodic orbits is found to occur for maximum semimajor axes near 1.0020 AU. In addition, near-periodic tadpole orbits appear to be stable against perturbations by Jupiter and Venus for periods of at least 10,000 yr. The possibility that minor planets actually exist in such orbits is considered.

Weissman, P. R.↗
NASA NTRS1993

(abstract) The Impactor Flux in the Pluto-Charon System

The impactor flux on Pluto and Charon consists of long-period comets from the Oort cloud passing through the region, and intermediate-period comets evolving inward to (or outward from) short-period orbits. The source of the intermediate-period comets is likely an extended disk of remnant planetesimals beyond the orbit of Neptune, commonly referred to as the Kuiper belt, though the Oort cloud may be an additional source of some short-period comets. The Pluto-Charon system actually circulates within the Kuiper belt, and with an aphelion of 49.3 AU, moves through a region where remnant planetesimal orbits would likely be stable over the age of the solar system. The recent discovery of two objects at Kuiper belt distances, 1992 QB 1 and 1993 FW, and the fact that the orbit determined for 1992 QB 1 is likely of low eccentricity and low inclination, have provided strong observational evidence for the Kuiper belt. Weisman et al (1989) showed that, for the expected number of objects in the Kuiper belt necessary to provide the short-period comet flux, cometary cratering on Pluto and Charon is dominated by Kuiper belt comets. We will provide updated estimates of the numbers, based on an improved understanding of Oort cloud and Kuiper belt dynamics since 1989, and on improved estimates for the radii and masses for Pluto and Charon. Because Pluto and Charon actually penetrate the stable region of the Kuiper belt beyond 40 AU, expected cratering rates for the pair are substantially higher than for Triton, which is located at the inner edge of the Kuiper belt. However, the Pluto-Charon system is tidally evolved and this may have led to resurfacing events at various times in its history. Thus, detailed predictions of the integrated crater density or comparisons with Triton are not very meaningful without a knowledge of that history. Differences in the suspected compositions of Pluto and Charon may also lead to different surface rheologies and thus, different retention times for cratered surfaces.

Pluto↗
NASA NTRS1990

Stationary orbits of satellites of disk galaxies

The satellite of an S-galaxy will experience opposing dynamical-friction forces from the stars of the disk and the halo. If these forces are in balance, the satellite may travel in a stable, near-circular orbit whose radius, for a wide range of physical parameters, should be limited to a zone 1.2 to 1.4 times the disk radius, much as is observed. The idea is very simple. The dynamical friction acting on a small satellite, moving through a stellar galactic halo, makes this satellite slow down. On the other hand, a stellar disk, rotating faster than a satellite, makes it speed up. But the density distributions in radius for disk's and halo's stars in real flat galaxies are quite different (respectively, exponential and power-law). Moreover, the observational data show that the exponential profile for disk's surface density drops abruptly at some radius (r sub d). So it is natural to expect that a stationary orbit could be near the edge of a disk (where two effects are mutually compensated).

Polyachenko, Valerij L.↗
NASA NTRS2021

Radar Observations from the Haystack Ultrawideband Satellite Imaging Radar in 2019

The NASA Orbital Debris Program Office (ODPO) relies primarily on ground-based radar measurements to characterize the distribution of small debris, down to approximately 3 mm depending upon altitude and the sensor used, in low Earth orbit (LEO). Since the early 1990’s, the Massachusetts Institute of Technology (MIT) Lincoln Laboratory (LL) has been collecting radar measurements for the NASA ODPO under agreements with the U.S. Department of Defense. The Haystack Ultrawideband Satellite Imaging Radar (HUSIR) is the primary ground-based radar sensor used by the ODPO and provides data on orbital debris down to an approximate size of 5.5 mm below 1000km altitude using the NASA size estimation model (SEM). Since orbital debris of this size are a significant risk to both human and robotic missions in LEO, the sensitivity of this radar makes it a high-value sensor. The NASA ODPO radar measurements are conducted on a continual basis for monitoring and enabling modeling of the orbital debris environment over time. HUSIR observations from 2019 are the most recent snapshot of the environment that has been measured and analyzed to date. In recent years, HUSIR measurements indicated relatively stable populations for the orbital debris objects that it is able to detect. In 2019, several interesting events happened on-orbit, including the start of large constellation deployments into LEO, as well as the Indian anti-satellite test with Microsat-R (International Designator 2019-006A, U.S. Strategic Command Space Surveillance Network catalog number 43947). Due to these events, coupled with a general increase in the number of missions and participants launching missions in recent years, continual monitoring is necessary to determine the effects of this increased activity on the orbital debris environment. This paper will explore the results of the 2019 HUSIR radar measurements, including above-average flux measurements at lower LEO altitudes and the evolution of the flux during the time of observations.

James Murray↗
NASA NTRSDate not supplied

Radar Observations from the Haystack Ultrawideband Satellite Imaging Radar in 2019

The NASA Orbital Debris Program Office (ODPO) relies primarily on ground-based radar measurements to characterize the distribution of small debris, down to approximately 3 mm depending upon altitude and the sensor used, in low Earth orbit (LEO). Since the early 1990’s, the Massachusetts Institute of Technology (MIT) Lincoln Laboratory (LL) has been collecting radar measurements for the NASA ODPO under agreements with the U.S. Department of Defense. The Haystack Ultrawideband Satellite Imaging Radar (HUSIR) is the primary ground-based radar sensor used by the ODPO and provides data on orbital debris down to an approximate size of 5.5 mm below 1000km altitude using the NASA size estimation model (SEM). Since orbital debris of this size are a significant risk to both human and robotic missions in LEO, the sensitivity of this radar makes it a high-value sensor. The NASA ODPO radar measurements are conducted on a continual basis for monitoring and enabling modeling of the orbital debris environment over time. HUSIR observations from 2019 are the most recent snapshot of the environment that has been measured and analyzed to date. In recent years, HUSIR measurements indicated relatively stable populations for the orbital debris objects that it is able to detect. In 2019, several interesting events happened on-orbit, including the start of large constellation deployments into LEO, as well as the Indian anti-satellite test with Microsat-R (International Designator 2019-006A, U.S. Strategic Command Space Surveillance Network catalog number 43947). Due to these events, coupled with a general increase in the number of missions and participants launching missions in recent years, continual monitoring is necessary to determine the effects of this increased activity on the orbital debris environment. This paper will explore the results of the 2019 HUSIR radar measurements, including above-average flux measurements at lower LEO altitudes and the evolution of the flux during the time of observations.

James Murray↗
NASA NTRSDate not supplied

Langley Automated Sensor Inter-calibration System (LASICS): Open Access Tools for Satellite Imager Inter-Calibration

Satellite imager calibration teams are tasked with maintaining stable measurement records to facilitate reliable monitoring of geophysical parameters and ensure dependable input for weather forecast models. Satellite imagers are neither uniformly calibrated nor radiometrically scaled to a common reference. Consistent inter-calibration between various earth-orbiting satellite imager pairs is a critical step in the creation of seamless earth-scene reflectance data records over time for input to higher level algorithms that retrieve earth system climate-sensitive properties. Each imager inherently by virtue of its optics (and associated properties like Spectral Response etc.) will have a unique measurement of the same earth-scene reflected signal. A key part of this is the identification and prediction of events where the imager pairs from their respective earth-orbits view the same stable terrestrial targets with nearly identical viewing and solar geometry. Langley Automated Sensor Inter-calibration System (LASICS) will provide the Earth remote sensing community with a foundation for the harmonization of these remote sensing records by the development of an intuitive, user-friendly, interactive web-based open access service that will leverage, an existing extensive and ever-growing database of earth-imager channel spectral response functions, on-demand capabilities for computation and visualization of spectral band adjustment factors using a variety of external hyper-spectral earth observation sources (e.g. SCIAMACHY, GOME-2 for VIS and IASI, AIRS for IR) and a variety of Solar Irradiance Spectra.

Arun Gopalan↗
NASA NTRS2014

SLS-Derived Lab: Precursor to Deep Space Human Exploration

Plans to send humans to Mars are in work and the launch system is being built. Are we ready? Robotic missions have successfully demonstrated transportation, entry, landing and surface operations but for human missions there are significant, potentially show-stopping issues. These issues, called Strategic Knowledge Gaps (SKGs) are the unanswered questions concerning long-duration exploration beyond low-earth-orbit. The gaps represent a risk of loss of life or mission and because they require extended exposure to the weightless environment outside earth's protective geo-magnetic field they cannot be resolved on the earth or on the International Space Station (ISS). Placing a laboratory at the relatively close and stable lunar Distant Retrograde Orbit (DRO) provides an accessible location with the requisite environmental conditions for conducting SKG research and testing mitigation solutions. Configurations comprised of multiple 3 meter and 4.3 meter diameter modules have been studied but the most attractive solution uses elements of the human Mars launch vehicle or Space Launch System (SLS) for a Mars proving ground laboratory. A shortened version of an SLS hydrogen propellant tank creates a Skylab-like pressure vessel that flies fully outfitted on a single launch. This not only offers significant savings by incorporating SLS pressure vessel development costs but avoids the expensive ISS approach using many launches with substantial on-orbit assembly before becoming operational. One of the most challenging SKGs is crew radiation protection; this is why SKG laboratory research is combined with Mars transit Habitat systems development. Fundamentally, the two cannot be divorced because using the habitat systems for protection requires actual hardware geometry and material properties intended to contribute to shielding effectiveness. The SKGs are difficult problems, solutions are not obvious, and require integrated, iterative, and multi-disciplinary development. A lunar DRO lab built from the launch system elements enables an early and representative transit habitat test bed necessary for closing gaps before sending humans on a 1000 day Mars mission.

Griffin, Brand↗
NASA NTRS2015

SLS-Derived Lab- Precursor to Deep Space Human Exploration

Plans to send humans to Mars are in the works and the launch system is being built. Are we ready? Transportation, entry, landing, and surface operations have been successfully demonstrated for robotic missions. However, for human missions, there are significant, potentially show-stopping issues. These issues, called Strategic Knowledge Gaps (SKGs), are the unanswered questions concerning long duration exploration Beyond low Earth Orbit (BEO). The gaps represent a risk of loss of life or mission and because they require extended exposure to the weightless environment outside of earth's protective geo-magnetic field, they cannot be resolved on Earth or on the International Space Station (ISS). Placing a laboratory at a relatively close and stable lunar Distant Retrograde Orbit (DRO) provides an accessible location with the requisite environmental conditions for conducting SKG research and testing mitigation solutions. Configurations comprised of multiple 3 m and 4.3 m diameter modules have been studied but the most attractive solution uses elements of the human Mars launch vehicle or Space Launch System (SLS) for a Mars proving ground laboratory. A shortened version of an SLS hydrogen propellant tank creates a Skylab-like pressure vessel that flies fully outfitted on a single launch. This not only offers significant savings by incorporating SLS pressure vessel development costs but avoids the expensive ISS approach using many launches with substantial on-orbit assembly before becoming operational. One of the most challenging SKGs is crew radiation protection; this is why SKG laboratory research is combined with Mars transit habitat systems development. Fundamentally, the two cannot be divorced because using the habitat systems for protection requires actual hardware geometry and material properties intended to contribute to shielding effectiveness. The SKGs are difficult problems. The solutions to these problems are not obvious; they require integrated, iterative, and multi-disciplinary development. A lunar DRO lab built from SLS elements enables an early and representative transit habitat test bed necessary for closing gaps before sending humans on a 1,000-day Mars mission.

Griffin, Brand M.↗
NASA NTRS2017

Human Cargo Resupply Logistics at Mars Using 150kW SEP Tug Cyclers

The ultimate goal of the Evolvable Mars Campaign is to build up a sustainable outpost at Mars that would be continually staffed with rotating crews. During this stage of human Mars exploration, it would be necessary to provision the crews with equipment and supplies both before and during their missions. In this paper, we study the use of 150 kW reusable SEP tugs as a means to deliver elements both to orbit and to the surface. The SEP tugs would make use of technology currently being developed for the proposed Asteroid Robotic Redirect Mission (ARRM). They would also be used to deliver food and supplies to sustain the crews similar to resupply missions for the International Space Station. The SEP tugs envisioned would be staged at a quasi-stable Lunar Near Rectilinear Orbit (NRO). The tugs would then mate with cargo vessels and xenon propellant being delivered by an SLS launch vehicle and continue on to Mars orbit where the cargo is delivered and the SEP tug returns to NRO to repeat the process. It was found that it is more efficient to deliver surface cargo via direct launch and entry versus using the tug cycler. Thousands of optimized low-thrust trajectories were simulated in order to create “Bacon plots” (like porkchop plots, but for low thrust transfers) in order to map out potential trajectories for dates from 2039 to 2052. This study maps out the buildup of a surface outpost as well as the necessary orbital and surface resupply launches in order to maintain it. In the steady state, a cadence of 9 cargo launches is required every 4 years to sustain the human outpost.

Post, Kevin↗