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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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The numerical accuracy of a satellite orbit computation program was investigated. The effects of roundoff and truncation errors on the solution were determined. New formulations that have the mean motion based on the total energy are compared to the more classical formulations and evaluated via numerical experiments.
Servicing economics of low Earth orbit satellites were studied. The following topics are examined: the economic importance of the repair missions; comparison of mission cost as opposed to satellite modulation transfer functions over a 10 year period; the effect of satellite flight rate change due to changes in satellite failure rate; estimated satellite cost reduction with shuttle operation projects from the 1960's to the 1970's; design objectives of the multimission modular spacecraft; and the economic importance of the repair mission.
Long term time-variation of osculating elements of satellite orbits perturbed by third body, including graphs of Kepler element variations along different orbits
Motion of near lunar satellite in initially circular orbit analyzed by integrating equations for time rates of change of orbit elements
A study of electromagnetic satellite de-spin system for orbiting satellites was conducted. Exact closed form solutions for the interacting magnetic flux have been formulated both for the case of the current loop and the bar magnet. The magnet system has been extended to the case of a U shaped magnet. Magnet design details are outlined. Computer simulation of the interacting flux density profile has been attempted. Data for the de-spin times has been extended to higher rpms.
A SPICE module for the Satellite Orbit Analysis Program (SOAP) precisely represents complex motion and maneuvers in an interactive, 3D animated environment with support for user-defined quantitative outputs. (SPICE stands for Spacecraft, Planet, Instrument, Camera-matrix, and Events). This module enables the SOAP software to exploit NASA mission ephemeris represented in the JPL Ancillary Information Facility (NAIF) SPICE formats. Ephemeris types supported include position, velocity, and orientation for spacecraft and planetary bodies including the Sun, planets, natural satellites, comets, and asteroids. Entire missions can now be imported into SOAP for 3D visualization, playback, and analysis. The SOAP analysis and display features can now leverage detailed mission files to offer the analyst both a numerically correct and aesthetically pleasing combination of results that can be varied to study many hypothetical scenarios. The software provides a modeling and simulation environment that can encompass a broad variety of problems using orbital prediction. For example, ground coverage analysis, communications analysis, power and thermal analysis, and 3D visualization that provide the user with insight into complex geometric relations are included. The SOAP SPICE module allows distributed science and engineering teams to share common mission models of known pedigree, which greatly reduces duplication of effort and the potential for error. The use of the software spans all phases of the space system lifecycle, from the study of future concepts to operations and anomaly analysis. It allows SOAP software to correctly position and orient all of the principal bodies of the Solar System within a single simulation session along with multiple spacecraft trajectories and the orientation of mission payloads. In addition to the 3D visualization, the user can define numeric variables and x-y plots to quantitatively assess metrics of interest.
The Engineering Orbit Analysis Unit at GE Valley Forge used an Intel Hypercube Parallel Processor to investigate the performance and gain experience of parallel processors with a multi-satellite orbit determination problem. A general study was selected in which major blocks of computation for the multi-satellite orbit computations were used as units to be assigned to the various processors on the Hypercube. Problems encountered or successes achieved in addressing the orbit determination problem would be more likely to be transferable to other parallel processors. The prime objective was to study the algorithm to allow processing of observations later in time than those employed in the state update. Expertise in ephemeris determination was exploited in addressing these problems and the facility used to bring a realism to the study which would highlight the problems which may not otherwise be anticipated. Secondary objectives were to gain experience of a non-trivial problem in a parallel processor environment, to explore the necessary interplay of serial and parallel sections of the algorithm in terms of timing studies, to explore the granularity (coarse vs. fine grain) to discover the granularity limit above which there would be a risk of starvation where the majority of nodes would be idle or under the limit where the overhead associated with splitting the problem may require more work and communication time than is useful.
Differential corrections applied to Vinti accurate reference satellite orbit with inclusion of third zonal harmonic
The authors describe simulation and analytical techniques for predicting the interference durations and protection ratios for deep-space ground stations from satellites in circular orbits. The simulation approach employs the characteristics of the ground station receivers in addition to the satellite specifications to calculate the daily statistics of the interference duration. On the other hand, the analytical method employs the ground station specifications to statistically predict the protection ratios as a function of the number of interference satellites and their associated probabilities. The results presented here can be used for both planning the missions and designing satellite communication links. The results can be also be used by spectrum managers and operations personnel to minimize the risk of harmful interference from satellites in low earth orbit to the deep space frequency band.
Baseline length and orientation of closely co-orbiting satellites estimated by combined onboard sensor and ground tracking data
Long term evolution of close low eccentricity lunar satellite orbits, describing lunar gravity effects by spherical harmonic expansion
The History of On-Orbit Satellite Fragmentations chronicles all known satellite fragmentation events, this 15th edition complete through a suspense date of 4 July 2018. Since the 14th edition breakups, in addition to launch activity, have resulted in an approximately 36% increase in the number of cataloged space objects. More significantly, breakup and anomalous debris accounted for 65% of the catalog growth observed. The reason for these large increases was the first accidental collision of two intact spacecraft, Iridium 33 and Cosmos 2251, and the continued cataloging of debris created by the intentional destruction of the Fengyun 1C spacecraft-the most environmentally harmful fragmentation to date.
To plan and execute a successful on-orbit satellite servicing mission, space support hardware, a serviceable spacecraft design, and flexible ground operations are required. This paper addresses space support hardware and in particular the Goddard Space Flight Center (GSFC) Satellite Servicing Project's (SSP) Flight Support System (FSS). The FSS is designed to be compatible with the Space Transportation System Orbiter and to support on-orbit servicing of Multimission Modular Spacecraft. It provides mechanical, electrical, and thermal support to the spacecraft during the servicing mission. Standardization of spacecraft design, components, and interfaces is required to make a satellite serviceable. Spacecraft ground operation capabilities must accommodate servicing mission requirements. Space support tools and equipment, such as the Power Ratchet Tool, assists the astronaut during time critical EVA operations to more efficiently perform the servicing mission.
Quasi-linearization techniques for solving boundary value satellite orbital transfer problems
Altimeter satellites are often maintained in a repeating orbit to facilitate the separation of sea-height variations from the geoid. However, atmospheric drag and solar radiation pressure cause a satellite orbit to drift. For Geosat this drift causes the ground track to vary by + or - 1 km about the nominal repeat path. This misalignment leads to an error in the estimates of sea surface height variations because of the local slope in the geoid. This error has been estimated globally for the Geosat Exact Repeat Mission using a mean sea surface constructed from Geos 3 and Seasat altimeter data. Over most of the ocean the geoid gradient is small, and the repeat-track misalignment leads to errors of only 1 to 2 cm. However, in the vicinity of trenches, continental shelves, islands, and seamounts, errors can exceed 20 cm. The estimated error is compared with direct estimates from Geosat altimetry, and a strong correlation is found in the vicinity of the Tonga and Aleutian trenches. This correlation increases as the orbit error is reduced because of the increased signal-to-noise ratio.
The targeting techniques are developed which are required to determine the guidance reference release time of the space shuttle navigation system, the orbital insertion targeting values, and a time line of orbital maneuvers. An extension is made for rendezvous with a target satellite in an elliptical orbit.
VIP program for computing incident and absorbed energy at orbiting satellite surface