Statistical initial orbit determination
NASA Astrophysics Data System (ADS)
Taff, L. G.; Belkin, B.; Schweiter, G. A.; Sommar, K.
1992-02-01
For the ballistic missile initial orbit determination problem in particular, the concept of 'launch folders' is extended. This allows to decouple the observational data from the initial orbit determination problem per se. The observational data is only used to select among the possible orbital element sets in the group of folders. Monte Carlo simulations using up to 7200 orbital element sets are described. The results are compared to the true orbital element set and the one a good radar would have been able to produce if collocated with the optical sensor. The simplest version of the new method routinely outperforms the radar initial orbital element set by a factor of two in future miss distance. In addition, not only can a differentially corrected orbital element set be produced via this approach - after only two measurements of direction - but also an updated, meaningful, six-dimensional covariance array for it can be calculated. This technique represents a significant advance in initial orbit determination for this problem, and the concept can easily be extended to minor planets and artificial satellites.
Analysis of initial orbit determination accuracy
NASA Astrophysics Data System (ADS)
Vananti, Alessandro; Schildknecht, Thomas
The Astronomical Institute of the University of Bern (AIUB) is conducting several search campaigns for orbital debris. The debris objects are discovered during systematic survey observations. In general only a short observation arc, or tracklet, is available for most of these objects. From this discovery tracklet a first orbit determination is computed in order to be able to find the object again in subsequent follow-up observations. The additional observations are used in the orbit improvement process to obtain accurate orbits to be included in a catalogue. In this paper, the accuracy of the initial orbit determination is analyzed. This depends on a number of factors: tracklet length, number of observations, type of orbit, astrometric error, and observation geometry. The latter is characterized by both the position of the object along its orbit and the location of the observing station. Different positions involve different distances from the target object and a different observing angle with respect to its orbital plane and trajectory. The present analysis aims at optimizing the geometry of the discovery observations depending on the considered orbit.
Dealing with Uncertainties in Initial Orbit Determination
NASA Technical Reports Server (NTRS)
Armellin, Roberto; Di Lizia, Pierluigi; Zanetti, Renato
2015-01-01
A method to deal with uncertainties in initial orbit determination (IOD) is presented. This is based on the use of Taylor differential algebra (DA) to nonlinearly map the observation uncertainties from the observation space to the state space. When a minimum set of observations is available DA is used to expand the solution of the IOD problem in Taylor series with respect to measurement errors. When more observations are available high order inversion tools are exploited to obtain full state pseudo-observations at a common epoch. The mean and covariance of these pseudo-observations are nonlinearly computed by evaluating the expectation of high order Taylor polynomials. Finally, a linear scheme is employed to update the current knowledge of the orbit. Angles-only observations are considered and simplified Keplerian dynamics adopted to ease the explanation. Three test cases of orbit determination of artificial satellites in different orbital regimes are presented to discuss the feature and performances of the proposed methodology.
NASA Technical Reports Server (NTRS)
Axelrad, Penina; Speed, Eden; Leitner, Jesse A. (Technical Monitor)
2002-01-01
This report summarizes the efforts to date in processing GPS measurements in High Earth Orbit (HEO) applications by the Colorado Center for Astrodynamics Research (CCAR). Two specific projects were conducted; initialization of the orbit propagation software, GEODE, using nominal orbital elements for the IMEX orbit, and processing of actual and simulated GPS data from the AMSAT satellite using a Doppler-only batch filter. CCAR has investigated a number of approaches for initialization of the GEODE orbit estimator with little a priori information. This document describes a batch solution approach that uses pseudorange or Doppler measurements collected over an orbital arc to compute an epoch state estimate. The algorithm is based on limited orbital element knowledge from which a coarse estimate of satellite position and velocity can be determined and used to initialize GEODE. This algorithm assumes knowledge of nominal orbital elements, (a, e, i, omega, omega) and uses a search on time of perigee passage (tau(sub p)) to estimate the host satellite position within the orbit and the approximate receiver clock bias. Results of the method are shown for a simulation including large orbital uncertainties and measurement errors. In addition, CCAR has attempted to process GPS data from the AMSAT satellite to obtain an initial estimation of the orbit. Limited GPS data have been received to date, with few satellites tracked and no computed point solutions. Unknown variables in the received data have made computations of a precise orbit using the recovered pseudorange difficult. This document describes the Doppler-only batch approach used to compute the AMSAT orbit. Both actual flight data from AMSAT, and simulated data generated using the Satellite Tool Kit and Goddard Space Flight Center's Flight Simulator, were processed. Results for each case and conclusion are presented.
Observability analysis for tracklet association and initial orbit determination
NASA Astrophysics Data System (ADS)
Siminski, Jan; Fiedler, Hauke
The geostationary orbit must be monitored to avoid accident-prone proximities of active satellites with space debris or other uncontrollable objects. Therefore, it should be scanned regularly by optical telescopes. Typical survey strategies divide the geostationary ring into right ascension and declination slots. Due to limited resources, each slot can only be observed for a short duration. The resulting measurement arcs, called tracklets, do not provide enough information to determine the full state of the object. Thus, the tracklets are associated to already known objects or combined with other measurements. The latter problem arises primarily in the catalog build-up phase, as well as when objects are lost and re-observed, e.g. if an object has been maneuvered and therefore cannot be successfully associated to an already cataloged object. The paper outlines a method that determines the orbit using the available information of two tracklets, i.e. their line-of-sights and their derivatives. The association and orbit determination is formulated as a boundary-value problem and solved using optimization schemes. The method uses the available information optimally, but fails to unambiguously associate closely spaced objects if the uncertainty attached to the line-of-sight derivative is too large. The difficulty increases with a larger time separation between the measurements. Due to unsuitable weather conditions, orbital slots might only be re-observed after one or more days. While a tracklet duration of 1-2 minutes provides enough information on the line-of-sight derivative for an association with a measurement of the same night, it cannot eliminate false association when the next measurement is taken in the following nights. To find a possible solution, a case study approach is used to analyze the association performance dependent on the tracklet duration and observation geometry. Observability conditions and their implications for consistent catalog maintenance are discussed. The results of this study can be used to improve current surveying strategies.
Coarse Initial Orbit Determination for a Geostationary Satellite Using Single-Epoch GPS Measurements
Kim, Ghangho; Kim, Chongwon; Kee, Changdon
2015-01-01
A practical algorithm is proposed for determining the orbit of a geostationary orbit (GEO) satellite using single-epoch measurements from a Global Positioning System (GPS) receiver under the sparse visibility of the GPS satellites. The algorithm uses three components of a state vector to determine the satelliteâ€™s state, even when it is impossible to apply the classical single-point solutions (SPS). Through consideration of the characteristics of the GEO orbital elements and GPS measurements, the components of the state vector are reduced to three. However, the algorithm remains sufficiently accurate for a GEO satellite. The developed algorithm was tested on simulated measurements from two or three GPS satellites, and the calculated maximum position error was found to be less than approximately 40 km or even several kilometers within the geometric range, even when the classical SPS solution was unattainable. In addition, extended Kalman filter (EKF) tests of a GEO satellite with the estimated initial state were performed to validate the algorithm. In the EKF, a reliable dynamic model was adapted to reduce the probability of divergence that can be caused by large errors in the initial state. PMID:25835299
NASA Astrophysics Data System (ADS)
Azimov, D.
The proposed approach aims to develop a new method of forming and processing of multiple hypotheses for initial orbit determination using optical observations. This method allows us to generalize the existing 2-dimensional flat constrained admissible region (CAR) to a unique 3-dimensional (3D) manifold of points corresponding to the pairs of observed right ascension and declination. Another advantage of this method is that unlike the existing methods of initial orbit determination using CAR, the range, range rate and angular rates are computed analytically using the angle observations, the location coordinates of the observation station, the semi-parameter and semi-major axis corresponding to the CAR. Unlike the existing 2D CAR, the 3D manifold does not include the pairs of range and range rate that do not correspond to the observed angles and computed range rat and angular rates. Given the the semi-parameter and semi-major axis, the proposed approach allows us to analytically compute the orientation angles as the Keplerian orbital elements, including the longitude of ascending node, inclination and argument of perigee. The resulting method represents a new and computationally efficient procedure for multiple data association processing through multiple hypotheses filter and allows for an uncertainty quantification.
NASA Astrophysics Data System (ADS)
Sease, Brad; Murphy, Timothy; Flewelling, Brien; Holzinger, Marcus J.; Black, Jonathan
2015-05-01
This paper presents an automatic RSO detection and tracking scheme operating at the optical sensor system level. The software presented is a pipeline for processing ground or space-based imagery built from several subalgorithms which processes raw or calibrated imagery, detects and discriminates non-star objects, and associates observations over time. An orbit determination routine uses an admissible region to start off an unscented particle filter. This preliminary orbit estimate allows prediction of the appearance of the object in the next frame. A matched filter uses this imagery to provide feedback to the initial detection and tracking process.
NASA Astrophysics Data System (ADS)
Kelecy, Tom; Shoemaker, Michael; Jah, Moriba
2013-08-01
A break-up in Low Earth Orbit (LEO) is simulated for 10 objects having area-to-mass ratios (AMR's) ranging from 0.1-10.0 m2/kg. The Constrained Admissible Region Multiple Hypothesis Filter (CAR-MHF) is applied to determining and characterizing the orbit and atmospheric drag parameters (CdA/m) simultaneously for each of the 10 objects with no a priori orbit or drag information. The results indicate that CAR-MHF shows promise for accurate, unambiguous and autonomous determination of the orbit and drag states.
NASA Technical Reports Server (NTRS)
Jordan, J. F.; Boggs, D. H.; Born, G. H.; Christensen, E. J.; Ferrari, A. J.; Green, D. W.; Hylkema, R. K.; Mohan, S. N.; Reinbold, S. J.; Sievers, G. L.
1973-01-01
A historic account of the activities of the Satellite OD Group during the MM'71 mission is given along with an assessment of the accuracy of the determined orbit of the Mariner 9 spacecraft. Preflight study results are reviewed, and the major error sources described. Tracking and data fitting strategy actually used in the real time operations is itemized, and Deep Space Network data available for orbit fitting during the mission and the auxiliary information used by the navigation team are described. A detailed orbit fitting history of the first four revolutions of the satellite orbit of Mariner 9 is presented, with emphasis on the convergence problems and the delivered solution for the first orbit trim maneuver. Also included are a solution accuracy summary, the history of the spacecraft orbit osculating elements, the results of verifying the radio solutions with TV imaging data, and a summary of the normal points generated for the relativity experiment.
Orbit determination in satellite geodesy
NASA Astrophysics Data System (ADS)
Beutler, G.; Schildknecht, T.; Hugentobler, U.; Gurtner, W.
2003-04-01
For centuries orbit determination in Celestial Mechanics was a synonym for the determination of six so-called Keplerian elements of the orbit of a minor planet or a comet based on a short series of (three or more) astrometric places observed from one or more observatories on the Earth's surface. With the advent of the space age the problem changed considerably in several respects: (1) orbits have to be determined for a new class of celestial objects, namely for artificial Earth satellites; (2) new observation types, in particular topocentric distances and radial velocities, are available for the establishment of highly accurate satellite orbits; (3) even for comparatively short arcs (up to a few revolutions) the orbit model that has to be used is much more complicated than for comparable problems in the planetary system: in addition to the gravitational perturbations due to Moon and planets higher-order terms in the Earth's gravity field have to be taken into account as well as non-gravitational effects like atmospheric drag and/or radiation pressure; (4) the parameter space is often of higher than the sixth dimension, because not only the six osculating elements referring to the initial epoch of an arc, but dynamical parameters defining the (a priori imperfectly known) force field have to be determined, as well. It may even be necessary to account for stochastic velocity changes. Orbit determination is not a well-known task in satellit geodesy. This is mainly due to the fact that orbit determination is often imbedded in a much more general parameter estimation problem, where other parameter types (referred to station positions, Earth rotation, atmosphere, etc.) have to be determined, as well. Three examples of "pure" orbit determination problems will be discussed subsequently: ? The first problem intends to optimize the observation process of one Satellite Laser Ranging (SLR) observatory. It is a filter problem, where the orbit is improved in real time with the goal to narrow down the so-called range-gate, defining the time interval when the echo of the LASER pulse is expected. ? Secondly we highlight orbit determination procedures (in particular advanced orbit parametrization techniques) related to the determination of the orbits of GPS satellites and of Low Earth Orbiters (LEOS) equipped with GPS receivers. ? We conclude by discussing the problem of determining the orbits of passive artificial satellites or of space debris using high-precision astrometric CCD-observations of these object.
ICESat Precision Orbit Determination
NASA Astrophysics Data System (ADS)
Rim, H.; Yoon, S.; Webb, C. E.; Kim, Y.; Schutz, B. E.
2003-12-01
Following the successful launch of the Ice, Cloud and land Elevation Satellite (ICESat) on January 13, 2003, 00:45 UTC, the GPS receiver on ICESat was turned on successfully on Jan. 17, 2003. High quality GPS data were collected since then to support Precision Orbit Determination (POD) activities. ICESat carries Geoscience Laser Altimeter System (GLAS) to measure ice-sheet topography and associated temporal changes, as well as cloud and atmospheric properties. To accomplish the ICESat science objectives, the position of the GLAS instrument in space should be determined with an accuracy of 5 cm and 20 cm in radial and horizontal components, respectively. This knowledge is acquired by the POD activities using the data collected by the GPS receiver on ICESat and the ground-based satellite laser ranging (SLR) data. It has been shown from pre-launch POD studies that the gravity model error is the dominant source of ICESat orbit errors. The predicted radial orbit errors at the ICESat orbit (600 km altitude) based on pre-launch gravity models, such as TEG-4 and EGM-96, are 7-15 cm. Performance of these gravity models and the recent gravity models from GRACE on ICESat POD were evaluated. The radial orbit accuracy is approaching 1-2 cm level with the GRACE gravity model. This paper also summarizes POD activities at Center for Space Research (CSR), which is responsible to generate ICESat POD products.
Transfer orbit determination accuracy for orbit maneuvers
NASA Astrophysics Data System (ADS)
Pinheiro, Mery Passos
This work intends to show the accuracy of the orbital elements determined during transfer orbit as a function of data span, as well as the feasibility of performance maneuvers. The orbit estimator used is a weighted least squares algorithm. The observation vector is composed of angle data (azimuth and elevation) and range data and are from the Astra IC mission. The state vector is either propagated by Brower model or numerical integration (for small eccentricities and inclination). The complete software to determine the orbit has been developed by Hughes Aircraft and been used for all Hughes satellite mission.
Lunar Reconnaissance Orbiter Orbit Determination Accuracy Analysis
NASA Technical Reports Server (NTRS)
Slojkowski, Steven E.
2014-01-01
Results from operational OD produced by the NASA Goddard Flight Dynamics Facility for the LRO nominal and extended mission are presented. During the LRO nominal mission, when LRO flew in a low circular orbit, orbit determination requirements were met nearly 100% of the time. When the extended mission began, LRO returned to a more elliptical frozen orbit where gravity and other modeling errors caused numerous violations of mission accuracy requirements. Prediction accuracy is particularly challenged during periods when LRO is in full-Sun. A series of improvements to LRO orbit determination are presented, including implementation of new lunar gravity models, improved spacecraft solar radiation pressure modeling using a dynamic multi-plate area model, a shorter orbit determination arc length, and a constrained plane method for estimation. The analysis presented in this paper shows that updated lunar gravity models improved accuracy in the frozen orbit, and a multiplate dynamic area model improves prediction accuracy during full-Sun orbit periods. Implementation of a 36-hour tracking data arc and plane constraints during edge-on orbit geometry also provide benefits. A comparison of the operational solutions to precision orbit determination solutions shows agreement on a 100- to 250-meter level in definitive accuracy.
NASA Technical Reports Server (NTRS)
Carpenter, James R.; Berry, Kevin; Gregpru. Late; Speckman, Keith; Hur-Diaz, Sun; Surka, Derek; Gaylor, Dave
2010-01-01
The Orbit Determination Toolbox is an orbit determination (OD) analysis tool based on MATLAB and Java that provides a flexible way to do early mission analysis. The toolbox is primarily intended for advanced mission analysis such as might be performed in concept exploration, proposal, early design phase, or rapid design center environments. The emphasis is on flexibility, but it has enough fidelity to produce credible results. Insight into all flight dynamics source code is provided. MATLAB is the primary user interface and is used for piecing together measurement and dynamic models. The Java Astrodynamics Toolbox is used as an engine for things that might be slow or inefficient in MATLAB, such as high-fidelity trajectory propagation, lunar and planetary ephemeris look-ups, precession, nutation, polar motion calculations, ephemeris file parsing, and the like. The primary analysis functions are sequential filter/smoother and batch least-squares commands that incorporate Monte-Carlo data simulation, linear covariance analysis, measurement processing, and plotting capabilities at the generic level. These functions have a user interface that is based on that of the MATLAB ODE suite. To perform a specific analysis, users write MATLAB functions that implement truth and design system models. The user provides his or her models as inputs to the filter commands. The software provides a capability to publish and subscribe to a software bus that is compliant with the NASA Goddard Mission Services Evolution Center (GMSEC) standards, to exchange data with other flight dynamics tools to simplify the flight dynamics design cycle. Using the publish and subscribe approach allows for analysts in a rapid design center environment to seamlessly incorporate changes in spacecraft and mission design into navigation analysis and vice versa.
Shadowing Lemma and chaotic orbit determination
NASA Astrophysics Data System (ADS)
Spoto, Federica; Milani, Andrea
2016-03-01
Orbit determination is possible for a chaotic orbit of a dynamical system, given a finite set of observations, provided the initial conditions are at the central time. The Shadowing Lemma (Anosov 1967; Bowen in J Differ Equ 18:333-356, 1975) can be seen as a way to connect the orbit obtained using the observations with a real trajectory. An orbit is a shadowing of the trajectory if it stays close to the real trajectory for some amount of time. In a simple discrete model, the standard map, we tackle the problem of chaotic orbit determination when observations extend beyond the predictability horizon. If the orbit is hyperbolic, a shadowing orbit is computed by the least squares orbit determination. We test both the convergence of the orbit determination iterative procedure and the behaviour of the uncertainties as a function of the maximum number of map iterations observed. When the initial conditions belong to a chaotic orbit, the orbit determination is made impossible by numerical instability beyond a computability horizon, which can be approximately predicted by a simple formula. Moreover, the uncertainty of the results is sharply increased if a dynamical parameter is added to the initial conditions as parameter to be estimated. The Shadowing Lemma does not dictate what the asymptotic behaviour of the uncertainties should be. These phenomena have significant implications, which remain to be studied, in practical problems of orbit determination involving chaos, such as the chaotic rotation state of a celestial body and a chaotic orbit of a planet-crossing asteroid undergoing many close approaches.
Shadowing Lemma and chaotic orbit determination
NASA Astrophysics Data System (ADS)
Spoto, Federica; Milani, Andrea
2015-12-01
Orbit determination is possible for a chaotic orbit of a dynamical system, given a finite set of observations, provided the initial conditions are at the central time. The Shadowing Lemma (Anosov 1967; Bowen in J Differ Equ 18:333-356, 1975) can be seen as a way to connect the orbit obtained using the observations with a real trajectory. An orbit is a shadowing of the trajectory if it stays close to the real trajectory for some amount of time. In a simple discrete model, the standard map, we tackle the problem of chaotic orbit determination when observations extend beyond the predictability horizon. If the orbit is hyperbolic, a shadowing orbit is computed by the least squares orbit determination. We test both the convergence of the orbit determination iterative procedure and the behaviour of the uncertainties as a function of the maximum number of map iterations observed. When the initial conditions belong to a chaotic orbit, the orbit determination is made impossible by numerical instability beyond a computability horizon, which can be approximately predicted by a simple formula. Moreover, the uncertainty of the results is sharply increased if a dynamical parameter is added to the initial conditions as parameter to be estimated. The Shadowing Lemma does not dictate what the asymptotic behaviour of the uncertainties should be. These phenomena have significant implications, which remain to be studied, in practical problems of orbit determination involving chaos, such as the chaotic rotation state of a celestial body and a chaotic orbit of a planet-crossing asteroid undergoing many close approaches.
Kaguya Orbit Determination from JPL
NASA Technical Reports Server (NTRS)
Haw, Robert J.; Mottinger, N. A.; Graat, E. J.; Jefferson, D. C.; Park, R.; Menom, P.; Higa, E.
2008-01-01
Selene (re-named 'Kaguya' after launch) is an unmanned mission to the Moon navigated, in part, by JPL personnel. Launched by an H-IIA rocket on September 14, 2007 from Tanegashima Space Center, Kaguya entered a high, Earth-centered phasing orbit with apogee near the radius of the Moon's orbit. After 19 days and two orbits of Earth, Kaguya entered lunar orbit. Over the next 2 weeks the spacecraft decreased its apolune altitude until reaching a circular, 100 kilometer altitude orbit. This paper describes NASA/JPL's participation in the JAXA/Kaguya mission during that 5 week period, wherein JPL provided tracking data and orbit determination support for Kaguya.
Orbit Determination of the Lunar Reconnaissance Orbiter
NASA Technical Reports Server (NTRS)
Mazarico, Erwan; Rowlands, D. D.; Neumann, G. A.; Smith, D. E.; Torrence, M. H.; Lemoine, F. G.; Zuber, M. T.
2011-01-01
We present the results on precision orbit determination from the radio science investigation of the Lunar Reconnaissance Orbiter (LRO) spacecraft. We describe the data, modeling and methods used to achieve position knowledge several times better than the required 50-100m (in total position), over the period from 13 July 2009 to 31 January 2011. In addition to the near-continuous radiometric tracking data, we include altimetric data from the Lunar Orbiter Laser Altimeter (LOLA) in the form of crossover measurements, and show that they strongly improve the accuracy of the orbit reconstruction (total position overlap differences decrease from approx.70m to approx.23 m). To refine the spacecraft trajectory further, we develop a lunar gravity field by combining the newly acquired LRO data with the historical data. The reprocessing of the spacecraft trajectory with that model shows significantly increased accuracy (approx.20m with only the radiometric data, and approx.14m with the addition of the altimetric crossovers). LOLA topographic maps and calibration data from the Lunar Reconnaissance Orbiter Camera were used to supplement the results of the overlap analysis and demonstrate the trajectory accuracy.
NASA Astrophysics Data System (ADS)
Stewart, William; Pratt, Alex R.; Entwisle, Leonard
2013-06-01
An overview is provided and first results presented from NEMETODE, The Network for Meteor Triangulation and Orbit Determination. This is a network of four low-light video cameras based in the North of England in the United Kingdom that use UFOCapture, UFOAnalyser and UFOOrbit to capture and analyse meteor data. NEMETODE is intended to supplement the increasing number of comparable teams around the world who are using similar networks. Many of these networks have been established to ascertain if the suspected meteor showers listed on the International Astronomical Union's Meteor Data Center actually exist and if so, determine if they can be associated with known parent bodies. This paper provides a detailed description of the equipment used and the techniques employed to collect and analyse the data. The results from the first full collaborative month of operation, 2012 August, are presented, with specific focus given to the 007 PER (Perseids) meteor shower. The Perseids are a well characterised shower and were selected to verify if the results from NEMETODE were consistent with currently accepted parameters.
Spitzer Orbit Determination During In-orbit Checkout Phase
NASA Technical Reports Server (NTRS)
Menon, Premkumar R.
2004-01-01
The Spitzer Space Telescope was injected into heliocentric orbit on August 25, 2003 to observe and study astrophysical phenomena in the infrared range of frequencies. The initial 60 days was dedicated to Spitzer's "In-Orbit Checkout (IOC)" efforts. During this time high levels of Helium venting were used to cool down the telescope. Attitude control was done using reaction wheels, which in turn were de-saturated using cold gas Nitrogen thrusting. Dense tracking data (nearly continuous) by the Deep Space network (DSN) were used to perform orbit determination and to assess any possible venting imbalance. Only Doppler data were available for navigation. This paper deals with navigation efforts during the IOC phase. It includes Dust Cover Ejection (DCE) monitoring, orbit determination strategy validation and results and assessment of non-gravitational accelerations acting on Spitzer including that due to possible imbalance in Helium venting.
Orbit Determination Issues for Libration Point Orbits
NASA Technical Reports Server (NTRS)
Beckman, Mark; Bauer, Frank (Technical Monitor)
2002-01-01
Libration point mission designers require knowledge of orbital accuracy for a variety of analyses including station keeping control strategies, transfer trajectory design, and formation and constellation control. Past publications have detailed orbit determination (OD) results from individual libration point missions. This paper collects both published and unpublished results from four previous libration point missions (ISEE (International Sun-Earth Explorer) -3, SOHO (Solar and Heliospheric Observatory), ACE (Advanced Composition Explorer) and MAP (Microwave Anisotropy Probe)) supported by Goddard Space Flight Center's Guidance, Navigation & Control Center. The results of those missions are presented along with OD issues specific to each mission. All past missions have been limited to ground based tracking through NASA ground sites using standard range and Doppler measurement types. Advanced technology is enabling other OD options including onboard navigation using seaboard attitude sensors and the use of the Very Long Baseline Interferometry (VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options potentially enable missions to reduce coherent dedicated tracking passes while maintaining orbital accuracy. With the increased projected loading of the DSN (Deep Space Network), missions must find alternatives to the standard OD scenario.
Shadowing Lemma and Chaotic Orbit Determination
NASA Astrophysics Data System (ADS)
Milani Comparetti, Andrea; Spoto, Federica
2015-08-01
Orbit determination is possible for a chaotic orbit of a dynamical system, given a finite set of observations, provided the initial conditions are at the central time. We test both the convergence of the orbit determination procedure and the behavior of the uncertainties as a function of the maximum number n of map iterations observed; this by using a simple discrete model, namely the standard map. Two problems appear: first, the orbit determination is made impossible by numerical instability beyond a computability horizon, which can be approximately predicted by a simple formula containing the Lyapounov time and the relative roundoff error. Second, the uncertainty of the results is sharply increased if a dynamical parameter (contained in the standard map formula) is added to the initial conditions as parameter to be estimated. In particular the uncertainty of the dynamical parameter, and of at least one of the initial conditions, decreases like n^a with a<0 but not large (of the order of unity). If only the initial conditions are estimated, their uncertainty decreases exponentially with n, thus it becomes very small. All these phenomena occur when the chosen initial conditions belong to a chaotic orbit (as shown by one of the well known Lyapounov indicators). If they belong to a non-chaotic orbit the computational horizon is much larger, if it exists at all, and the decrease of the uncertainty appears to be polynomial in all parameters, like n^a with a approximately 1/2; the difference between the case with and without dynamical parameter estimated disappears. These phenomena, which we can investigate in a simple model, have significant implications in practical problems of orbit determination involving chatic phenomena, such as the chaotic rotation state of a celestial body and a chaotic orbit of a planet-crossing asteroid undergoing many close approaches.
Lunar Reconnaissance Orbiter Orbit Determination Accuracy Analysis
NASA Technical Reports Server (NTRS)
Slojkowski, Steven E.
2014-01-01
LRO definitive and predictive accuracy requirements were easily met in the nominal mission orbit, using the LP150Q lunar gravity model. center dot Accuracy of the LP150Q model is poorer in the extended mission elliptical orbit. center dot Later lunar gravity models, in particular GSFC-GRAIL-270, improve OD accuracy in the extended mission. center dot Implementation of a constrained plane when the orbit is within 45 degrees of the Earth-Moon line improves cross-track accuracy. center dot Prediction accuracy is still challenged during full-Sun periods due to coarse spacecraft area modeling - Implementation of a multi-plate area model with definitive attitude input can eliminate prediction violations. - The FDF is evaluating using analytic and predicted attitude modeling to improve full-Sun prediction accuracy. center dot Comparison of FDF ephemeris file to high-precision ephemeris files provides gross confirmation that overlap compares properly assess orbit accuracy.
Gravity Probe B orbit determination
NASA Astrophysics Data System (ADS)
Shestople, P.; Ndili, A.; Hanuschak, G.; Parkinson, B. W.; Small, H.
2015-11-01
The Gravity Probe B (GP-B) satellite was equipped with a pair of redundant Global Positioning System (GPS) receivers used to provide navigation solutions for real-time and post-processed orbit determination (OD), as well as to establish the relation between vehicle time and coordinated universal time. The receivers performed better than the real-time position requirement of 100 m rms per axis. Post-processed solutions indicated an rms position error of 2.5 m and an rms velocity error of 2.2 mm s-1. Satellite laser ranging measurements provided independent verification of the GPS-derived GP-B orbit. We discuss the modifications and performance of the Trimble Advance Navigation System Vector III GPS receivers. We describe the GP-B precision orbit and detail the OD methodology, including ephemeris errors and the laser ranging measurements.
Low thrust orbit determination program
NASA Technical Reports Server (NTRS)
Hong, P. E.; Shults, G. L.; Huling, K. R.; Ratliff, C. W.
1972-01-01
Logical flow and guidelines are provided for the construction of a low thrust orbit determination computer program. The program, tentatively called FRACAS (filter response analysis for continuously accelerating spacecraft), is capable of generating a reference low thrust trajectory, performing a linear covariance analysis of guidance and navigation processes, and analyzing trajectory nonlinearities in Monte Carlo fashion. The choice of trajectory, guidance and navigation models has been made after extensive literature surveys and investigation of previous software. A key part of program design relied upon experience gained in developing and using Martin Marietta Aerospace programs: TOPSEP (Targeting/Optimization for Solar Electric Propulsion), GODSEP (Guidance and Orbit Determination for SEP) and SIMSEP (Simulation of SEP).
Mars Science Laboratory Orbit Determination
NASA Technical Reports Server (NTRS)
Kruizinga, Gerhard; Gustafson, Eric; Jefferson, David; Martin-Mur, Tomas; Mottinger, Neil; Pelletier, Fred; Ryne, Mark; Thompson, Paul
2012-01-01
Mars Science Laboratory (MSL) Orbit Determination (OD) met all requirements with considerable margin, MSL OD team developed spin signature removal tool and successfully used the tool during cruise, A novel approach was used for the MSL solar radiation pressure model and resulted in a very accurate model during the approach phase, The change in velocity for Attitude Control System (ACS) turns was successfully calibrated and with appropriate scale factor resulted in improved change in velocity prediction for future turns, All Trajectory Correction Maneuvers were successfully reconstructed and execution errors were well below the assumed pre-fight execution errors, The official OD solutions were statistically consistent throughout cruise and for OD solutions with different arc lengths as well, Only EPU-1 was sent to MSL. All other Entry Parameter Updates were waived, EPU-1 solution was only 200 m separated from final trajectory reconstruction in the B-plane
Information Measures for Statistical Orbit Determination
ERIC Educational Resources Information Center
Mashiku, Alinda K.
2013-01-01
The current Situational Space Awareness (SSA) is faced with a huge task of tracking the increasing number of space objects. The tracking of space objects requires frequent and accurate monitoring for orbit maintenance and collision avoidance using methods for statistical orbit determination. Statistical orbit determination enables us to obtain…
Information Measures for Statistical Orbit Determination
ERIC Educational Resources Information Center
Mashiku, Alinda K.
2013-01-01
The current Situational Space Awareness (SSA) is faced with a huge task of tracking the increasing number of space objects. The tracking of space objects requires frequent and accurate monitoring for orbit maintenance and collision avoidance using methods for statistical orbit determination. Statistical orbit determination enables us to obtainâ€¦
NASA Astrophysics Data System (ADS)
Shakun, L. S.; Koshkin, N. I.
2014-06-01
The number of artificial space objects in the low Earth orbit has been continuously increasing. That raises the requirements for the accuracy of measurement of their coordinates and for the precision of the prediction of their motion. The accuracy of the prediction can be improved if the actual current orientation of the non-spherical satellite is taken into account. In so doing, it becomes possible to directly determine the atmospheric density along the orbit. The problem solution is to regularly conduct the photometric surveillances of a large number of satellites and monitor the parameters of their rotation around the centre of mass. To do that, it is necessary to get and promptly process large video arrays, containing pictures of a satellite against the background stars. In the present paper, the method for the simultaneous measurement of coordinates and brightness of the low Earth orbit space objects against the background stars when they are tracked by telescope KT-50 with the mirror diameter of 50 cm and with video camera WAT-209H2 is considered. The problem of determination of the moments of exposures of images is examined in detail. The estimation of the accuracy of measuring both the apparent coordinates of stars and their photometry is given on the example of observation of the open star cluster. In the presented observations, the standard deviation of one position measured is 1?, the accuracy of determination of the moment of exposure of images is better than 0.0001 s. The estimate of the standard deviation of one measurement of brightness is 0.1m. Some examples of the results of surveillances of satellites are also presented in the paper.
Precision Orbit Determination for the Lunar Reconnaissance Orbiter
NASA Astrophysics Data System (ADS)
Lemoine, F. G.; Mazarico, E.; Rowlands, D. D.; Torrence, M. H.; McGarry, J. F.; Neumann, G. A.; Mao, D.; Smith, D. E.; Zuber, M. T.
2010-05-01
The Lunar Reconnaissance Orbiter (LRO) spacecraft was launched on June 18, 2009. In mid-September 2009, the spacecraft orbit was changed from its commissioning orbit (30 x 216 km polar) to a quasi-frozen polar orbit with an average altitude of 50km (+-15km). One of the goals of the LRO mission is to develop a new lunar reference frame to facilitate future exploration. Precision Orbit Determination is used to achieve the accuracy requirements, and to precisely geolocate the high-resolution datasets obtained by the LRO instruments. In addition to the tracking data most commonly used to determine spacecraft orbits in planetary missions (radiometric Range and Doppler), LRO benefits from two other types of orbital constraints, both enabled by the Lunar Orbiter Laser Altimeter (LOLA) instrument. The altimetric data collected as the instrument's primary purpose can be used to derive constraints on the orbit geometry at the times of laser groundtrack intersections (crossovers). The multi-beam configuration and high firing-rate of LOLA further improves the strength of these crossovers, compared to what was possible with the MOLA instrument onboard Mars Global Surveyor (MGS). Furthermore, one-way laser ranges (LR) between Earth International Laser Ranging Service (ILRS) stations and the spacecraft are made possible by the addition of a small telescope mounted on the spacecraft high-gain antenna. The photons received from Earth are transmitted to one LOLA detector by a fiber optics bundle. Thanks to the accuracy of the LOLA timing system, the precision of 5-s LR normal points is below 10cm. We present the first results of the Precision Orbit Determination (POD) of LRO through the commissioning and nominal phases of the mission. Orbit quality is discussed, and various gravity fields are evaluated with the new (independent) LRO radio tracking data. The altimetric crossovers are used as an independent data type to evaluate the quality of the orbits. The contribution of the LR data is assessed. Multi-arc solutions over entire months are presented, which allow to strengthen the LR data because fewer clock-related parameters need to be adjusted. Finally, a preliminary 1-month solution with altimetric crossover constraints is evaluated and discussed
Orbit determination by range-only data.
NASA Technical Reports Server (NTRS)
Duong, N.; Winn, C. B.
1973-01-01
The determination of satellite orbits for use in geodesy using range-only data has been examined. A recently developed recursive algorithm for rectification of the nominal orbit after processing each observation has been tested. It is shown that when a synchronous satellite is tracked simultaneously with a subsynchronous geodetic target satellite, the orbits of each may be readily determined by processing the range information. Random data errors and satellite perturbations are included in the examples presented.
Orbit Determination Analysis Utilizing Radiometric and Laser Ranging Measurements for GPS Orbit
NASA Technical Reports Server (NTRS)
Welch, Bryan W.
2007-01-01
While navigation systems for the determination of the orbit of the Global Position System (GPS) have proven to be very effective, the current issues involve lowering the error in the GPS satellite ephemerides below their current level. In this document, the results of an orbit determination covariance assessment are provided. The analysis is intended to be the baseline orbit determination study comparing the benefits of adding laser ranging measurements from various numbers of ground stations. Results are shown for two starting longitude assumptions of the satellite location and for nine initial covariance cases for the GPS satellite state vector.
Determination of GPS orbits to submeter accuracy
NASA Technical Reports Server (NTRS)
Bertiger, W. I.; Lichten, S. M.; Katsigris, E. C.
1988-01-01
Orbits for satellites of the Global Positioning System (GPS) were determined with submeter accuracy. Tests used to assess orbital accuracy include orbit comparisons from independent data sets, orbit prediction, ground baseline determination, and formal errors. One satellite tracked 8 hours each day shows rms error below 1 m even when predicted more than 3 days outside of a 1-week data arc. Differential tracking of the GPS satellites in high Earth orbit provides a powerful relative positioning capability, even when a relatively small continental U.S. fiducial tracking network is used with less than one-third of the full GPS constellation. To demonstrate this capability, baselines of up to 2000 km in North America were also determined with the GPS orbits. The 2000 km baselines show rms daily repeatability of 0.3 to 2 parts in 10 to the 8th power and agree with very long base interferometry (VLBI) solutions at the level of 1.5 parts in 10 to the 8th power. This GPS demonstration provides an opportunity to test different techniques for high-accuracy orbit determination for high Earth orbiters. The best GPS orbit strategies included data arcs of at least 1 week, process noise models for tropospheric fluctuations, estimation of GPS solar pressure coefficients, and combine processing of GPS carrier phase and pseudorange data. For data arc of 2 weeks, constrained process noise models for GPS dynamic parameters significantly improved the situation.
A new chapter in precise orbit determination
NASA Technical Reports Server (NTRS)
Yunck, T. P.
1992-01-01
A report is presented on the use of GPS receivers on board orbiting spacecraft to determine their orbits with unprecedented accuracy. By placing a GPS receiver aboard a satellite one can observe its true motion and reconstruct its trajectory in great detail without knowledge of the forces acting on it. Only the accuracy of the GPS carrier-phase observable, which can be better than 1 cm for a 1 sec duration observation, ultimately limits 'user orbit' accuracy.
Modeling issues in precision orbit determination for Mars orbiter
NASA Technical Reports Server (NTRS)
Lemoine, Frank G.; Rosborough, George W.; Smith, David E.
1990-01-01
This paper examines the accuracy of recent Mars gravity models and the importance of perturbations due to the Mars radiation pressure and the Martian moons, Phobos and Deimos, on the trajectories of Mars orbiters. A linear orbit perturbation theory is used to characterize the patterns of gravity field near resonances for the Viking and Mariner 9 spacecraft. These resonances are shown to have considerable power and their potential for contributing to Mars gravity solutions is emphasized. It is shown that some of the same resonance orders which appear in the Viking orbits, dominate the radial orbit error spectrum for Mars Observer. Results of orbit determination simulations at the Goddard Space Flight Center show that the perturbations caused by the Martian moons and the Mars radiation pressure are larger than 0.1 mm/s, the expected precision of the Mars Observer Doppler tracking data. Tests with the Viking Doppler data indicate that best analysis of these data mandates the inclusion of the Phobos gravitational perturbation in the modeling of Viking spacecraft trajectories.
Precision orbit determination of altimetric satellites
NASA Technical Reports Server (NTRS)
Shum, C. K.; Ries, John C.; Tapley, Byron D.
1994-01-01
The ability to determine accurate global sea level variations is important to both detection and understanding of changes in climate patterns. Sea level variability occurs over a wide spectrum of temporal and spatial scales, and precise global measurements are only recently possible with the advent of spaceborne satellite radar altimetry missions. One of the inherent requirements for accurate determination of absolute sea surface topography is that the altimetric satellite orbits be computed with sub-decimeter accuracy within a well defined terrestrial reference frame. SLR tracking in support of precision orbit determination of altimetric satellites is significant. Recent examples are the use of SLR as the primary tracking systems for TOPEX/Poseidon and for ERS-1 precision orbit determination. The current radial orbit accuracy for TOPEX/Poseidon is estimated to be around 3-4 cm, with geographically correlated orbit errors around 2 cm. The significance of the SLR tracking system is its ability to allow altimetric satellites to obtain absolute sea level measurements and thereby provide a link to other altimetry measurement systems for long-term sea level studies. SLR tracking allows the production of precise orbits which are well centered in an accurate terrestrial reference frame. With proper calibration of the radar altimeter, these precise orbits, along with the altimeter measurements, provide long term absolute sea level measurements. The U.S. Navy's Geosat mission is equipped with only Doppler beacons and lacks laser retroreflectors. However, its orbits, and even the Geosat orbits computed using the available full 40-station Tranet tracking network, yield orbits with significant north-south shifts with respect to the IERS terrestrial reference frame. The resulting Geosat sea surface topography will be tilted accordingly, making interpretation of long-term sea level variability studies difficult.
Orbit determination using synthetic aperture radar
NASA Technical Reports Server (NTRS)
Taber, W. L.; Synnott, S. P.; Riedel, J. E.
1986-01-01
The use of synthetic aperture radar (SAR) images to estimate orbital parameters is studied. The SAR image formation process which requires the ability to repeatedly transmit identical signals and accurately sense the return echoes from a region of terrain is described. The orbit determination capabilities of the SAR system's observables are investigated. Five SAR observations were collected from a simulated shuttle orbit, which was circular with a latitude of 220 km and along-track velocity of 7.7 km/sec, to obtain along-track and line-of-sight direction position measurements; the simulation reveals that only three SAR observations were required to determine the position of the spacecraft to within 100 m. A prototype SAR orbit determination system was developed. The system consists of a VAX 11/780 time-shared computer, a frame buffer, topographic maps, and software for line-pixel location of an object within a SAR image and for orbit determination. The prototype is applied to the processing of a single short arc of Shuttle Imaging-Radar-B (SIR-B) data. It is observed that the SAR data is useful as orbit determination or tracking data; however, the low SNRs in the SIR-B data made feature identification difficult.
NASA Technical Reports Server (NTRS)
Quast, Peter; Tung, Frank; West, Mark; Wider, John
2000-01-01
The Chandra X-ray Observatory (CXO, formerly AXAF) is the third of the four NASA great observatories. It was launched from Kennedy Space Flight Center on 23 July 1999 aboard the Space Shuttle Columbia and was successfully inserted in a 330 x 72,000 km orbit by the Inertial Upper Stage (IUS). Through a series of five Integral Propulsion System burns, CXO was placed in a 10,000 x 139,000 km orbit. After initial on-orbit checkout, Chandra's first light images were unveiled to the public on 26 August, 1999. The CXO Pointing Control and Aspect Determination (PCAD) subsystem is designed to perform attitude control and determination functions in support of transfer orbit operations and on-orbit science mission. After a brief description of the PCAD subsystem, the paper highlights the PCAD activities during the transfer orbit and initial on-orbit operations. These activities include: CXO/IUS separation, attitude and gyro bias estimation with earth sensor and sun sensor, attitude control and disturbance torque estimation for delta-v burns, momentum build-up due to gravity gradient and solar pressure, momentum unloading with thrusters, attitude initialization with star measurements, gyro alignment calibration, maneuvering and transition to normal pointing, and PCAD pointing and stability performance.
The GEOS-3 orbit determination investigation
NASA Technical Reports Server (NTRS)
Pisacane, V. L.; Eisner, A.; Yionoulis, S. M.; Mcconahy, R. J.; Black, H. D.; Pryor, L. L.
1978-01-01
The nature and improvement in satellite orbit determination when precise altimetric height data are used in combination with conventional tracking data was determined. A digital orbit determination program was developed that could singly or jointly use laser ranging, C-band ranging, Doppler range difference, and altimetric height data. Two intervals were selected and used in a preliminary evaluation of the altimeter data. With the data available, it was possible to determine the semimajor axis and eccentricity to within several kilometers, in addition to determining an altimeter height bias. When used jointly with a limited amount of either C-band or laser range data, it was shown that altimeter data can improve the orbit solution.
Orbit determination methods in view of the PODET project
NASA Astrophysics Data System (ADS)
Deleflie, F.; Coulot, D.; Decosta, R.; Richard, P.
2013-11-01
We present an orbit determination method based on genetic algorithms. Contrary to usual estimation methods mainly based on least-squares methods, these algorithms do not require any a priori knowledge of the initial state vector to be estimated. These algorithms can be applied when a new satellite is launched or for uncatalogued objects We show in this paper preliminary results obtained from an SLR satellite, for which tracking data acquired by the ILRS network enable to build accurate orbital arcs at a few centimeter level, which can be used as a reference orbit. The method is carried out in several steps: (i) an analytical propagation of the equations of motion, (ii) an estimation kernel based on genetic algorithms, which follows the usual steps of such approaches: initialization and evolution of a selected population, so as to determine the best parameters. Each parameter to be estimated, namely each initial keplerian element, has to be searched among an interval that is preliminary chosen.
Meteor orbit determination with improved accuracy
NASA Astrophysics Data System (ADS)
Dmitriev, Vasily; Lupovla, Valery; Gritsevich, Maria
2015-08-01
Modern observational techniques make it possible to retrive meteor trajectory and its velocity with high accuracy. There has been a rapid rise in high quality observational data accumulating yearly. This fact creates new challenges for solving the problem of meteor orbit determination. Currently, traditional technique based on including corrections to zenith distance and apparent velocity using well-known Schiaparelli formula is widely used. Alternative approach relies on meteoroid trajectory correction using numerical integration of equation of motion (Clark & Wiegert, 2011; Zuluaga et al., 2013). In our work we suggest technique of meteor orbit determination based on strict coordinate transformation and integration of differential equation of motion. We demonstrate advantage of this method in comparison with traditional technique. We provide results of calculations by different methods for real, recently occurred fireballs, as well as for simulated cases with a priori known retrieval parameters. Simulated data were used to demonstrate the condition, when application of more complex technique is necessary. It was found, that for several low velocity meteoroids application of traditional technique may lead to dramatically delusion of orbit precision (first of all, due to errors in ?, because this parameter has a highest potential accuracy). Our results are complemented by analysis of sources of perturbations allowing to quantitatively indicate which factors have to be considered in orbit determination. In addition, the developed method includes analysis of observational error propagation based on strict covariance transition, which is also presented.Acknowledgements. This work was carried out at MIIGAiK and supported by the Russian Science Foundation, project No. 14-22-00197.References:Clark, D. L., & Wiegert, P. A. (2011). A numerical comparison with the Ceplecha analytical meteoroid orbit determination method. Meteoritics & Planetary Science, 46(8), pp. 1217-1225.Zuluaga, J. I., et al. (2013). The orbit of the Chelyabinsk event impactor as reconstructed from amateur and public footage. Earth and Planetary Science Letters arXiv:1303.1796. Retrieved march 7, 2013
Automated GPS-based operational orbit determination
NASA Astrophysics Data System (ADS)
Meek, Matthew Cameron
Satellite operations depend on being able to generate accurate predictions of a spacecraft's orbit in a very short period of time, typically a few hours, after observations are made. The satellite ephemeris generated in this process is used by mission controllers for planning operations such as vehicle pointing and orbit adjust generation. The research described in this dissertation, investigates the methods and parameterizations necessary to achieve a fast and accurate ephemeris. To accomplish these investigations, an automated system is used. Two distinct spacecraft missions are discussed. They each have specific goals that must be met by their operational orbit determination systems. The first is ICESat, a scientific satellite that is part of NASA's Earth Observation System (EOS), and is operated by the Laboratory for Atmospheric and Space Physics (LASP). The primary OD requirement for ICESat is to provide predictions accurate to 10 meters cross-track for 48 hours to accomplish instrument pointing planning. The second mission is Quick-Bird, a commercial imaging satellite that is owned and operated by Digital Globe, Inc. QuickBird requires post-processed orbits with 3 meters (1sigma) accuracy in total position and 30 day orbit predictions to accomplish imagery planning. A variety of measurement processing schemes and error corrections are explored for each of these spacecraft. It is shown that it is possible to achieve approximately one meter (1sigma) orbits for both spacecraft in a orbit determination system that is designed for use in spacecraft operations. In the ICESat case, it was found that using single-differenced measurements met the requirements while reducing both the processing time and the logistical load for importing external data. QuickBird benefitted from the addition of the DRVID method of ionospheric removal and from using double-differenced measurements.
NASA Technical Reports Server (NTRS)
Yee, C. P.; Kelbel, D. A.; Lee, T.; Dunham, J. B.; Mistretta, G. D.
1990-01-01
The influence of ionospheric refraction on orbit determination was studied through the use of the Orbit Determination Error Analysis System (ODEAS). The results of a study of the orbital state estimate errors due to the ionospheric refraction corrections, particularly for measurements involving spacecraft-to-spacecraft tracking links, are presented. In current operational practice at the Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF), the ionospheric refraction effects on the tracking measurements are modeled in the Goddard Trajectory Determination System (GTDS) using the Bent ionospheric model. While GTDS has the capability of incorporating the ionospheric refraction effects for measurements involving ground-to-spacecraft tracking links, such as those generated by the Ground Spaceflight Tracking and Data Network (GSTDN), it does not have the capability to incorporate the refraction effects for spacecraft-to-spacecraft tracking links for measurements generated by the Tracking and Data Relay Satellite System (TDRSS). The lack of this particular capability in GTDS raised some concern about the achievable accuracy of the estimated orbit for certain classes of spacecraft missions that require high-precision orbits. Using an enhanced research version of GTDS, some efforts have already been made to assess the importance of the spacecraft-to-spacecraft ionospheric refraction corrections in an orbit determination process. While these studies were performed using simulated data or real tracking data in definitive orbit determination modes, the study results presented here were obtained by means of covariance analysis simulating the weighted least-squares method used in orbit determination.
Filtering theory applied to orbit determination
NASA Technical Reports Server (NTRS)
Torroglosa, V.
1973-01-01
Modifications to the extended Kalman filter and the Jazwinski filter are made and compared with the classical extended Kalman filter in applications to orbit determination using real data. The results show that with the kind of data available today, the application of filtering theories in this field presents many problems.
Autonomous landmark tracking orbit determination strategy
NASA Technical Reports Server (NTRS)
Miller, J. K.; Cheng, Y.
2003-01-01
In this paper, an orbit determination strategy is described that is fully autonomous and relies on a computer-based crater detection and identification algorithm that is suitable for both automation of the ground based navigation system and autonomous spacecraft based navigation.
Algorithms for Autonomous GS Orbit Determination and Formation Flying
NASA Technical Reports Server (NTRS)
Moreau, Michael C.; Speed, Eden Denton-Trost; Axelrad, Penina; Leitner, Jesse (Technical Monitor)
2001-01-01
This final report for our study of autonomous Global Positioning System (GPS) satellite orbit determination comprises two sections. The first is the Ph.D. dissertation written by Michael C. Moreau entitled, "GPS Receiver Architecture for Autonomous Navigation in High Earth Orbits." Dr. Moreau's work was conducted under both this project and a NASA GSRP. His dissertation describes the key design features of a receiver specifically designed for autonomous operation in high earth orbits (HEO). He focused on the implementation and testing of these features for the GSFC PiVoT receiver. The second part is a memo describing a robust method for autonomous initialization of the orbit estimate given very little a priori information and sparse measurements. This is a key piece missing in the design of receivers for HEO.
Statistical inversion method for binary asteroids orbit determination
NASA Astrophysics Data System (ADS)
Kovalenko, I.; Hestroffer, D.; Doressoundiram, A.; Emelyanov, N.; Stoica, R.
2015-08-01
We focus on the study of binary asteroids, which are common in the Solar system from its inner to its outer regions. These objects provide fundamental physical parameters such as mass and density, and hence clues about the early Solar System. The present method of orbit calculation for resolved binaries is based on Markov Chain Monte-Carlo statistical inversion technique. In particular, we use the Metropolis-Hastings algorithm combined the Thiele-Innes equation for sampling orbital elements through the sampling of observations. The method requires a minimum of four observations, made at the same tangent plane; it is of particular interest for initial orbit determination. The observations are sampled within their observational errors with an assumed distribution. The sampling predicts the whole region of possible orbits, including the one that is most probable.
James Webb Space Telescope Orbit Determination Analysis
NASA Technical Reports Server (NTRS)
Yoon, Sungpil; Rosales, Jose; Richon, Karen
2014-01-01
The James Webb Space Telescope (JWST) is designed to study and answer fundamental astrophysical questions from an orbit about the Sun-Earth/Moon L2 libration point, 1.5 million km away from Earth. This paper describes the results of an orbit determination (OD) analysis of the JWST mission emphasizing the challenges specific to this mission in various mission phases. Three mid-course correction (MCC) maneuvers during launch and early orbit phase and transfer orbit phase are required for the spacecraft to reach L2. These three MCC maneuvers are MCC-1a at Launch+12 hours, MCC-1b at L+2.5 days and MCC-2 at L+30 days. Accurate OD solutions are needed to support MCC maneuver planning. A preliminary analysis shows that OD performance with the given assumptions is adequate to support MCC maneuver planning. During the nominal science operations phase, the mission requires better than 2 cm/sec velocity estimation performance to support stationkeeping maneuver planning. The major challenge to accurate JWST OD during the nominal science phase results from the unusually large solar radiation pressure force acting on the huge sunshield. Other challenges are stationkeeping maneuvers at 21-day intervals to keep JWST in orbit around L2, frequent attitude reorientations to align the JWST telescope with its targets and frequent maneuvers to unload momentum accumulated in the reaction wheels. Monte Carlo analysis shows that the proposed OD approach can produce solutions that meet the mission requirements.
James Webb Space Telescope Orbit Determination Analysis
NASA Technical Reports Server (NTRS)
Yoon, Sungpil; Rosales, Jose; Richon, Karen
2014-01-01
The James Webb Space Telescope (JWST) is designed to study and answer fundamental astrophysical questions from an orbit about the Sun-EarthMoon L2 libration point, 1.5 million km away from Earth. Three mid-course correction (MCC) maneuvers during launch and early orbit phase and transfer orbit phase are required for the spacecraft to reach L2. These three MCC maneuvers are MCC-1a at Launch+12 hours, MCC-1b at L+2.5 days and MCC-2 at L+30 days. Accurate orbit determination (OD) solutions are needed to support MCC maneuver planning. A preliminary analysis shows that OD performance with the given assumptions is adequate to support MCC maneuver planning. During the nominal science operations phase, the mission requires better than 2 cmsec velocity estimation performance to support stationkeeping maneuver planning. The major challenge to accurate JWST OD during the nominal science phase results from the unusually large solar radiation pressure force acting on the huge sunshield. Other challenges are stationkeeping maneuvers at 21-day intervals to keep JWST in orbit around L2, frequent attitude reorientations to align the JWST telescope with its targets and frequent maneuvers to unload momentum accumulated in the reaction wheels. Monte Carlo analysis shows that the proposed OD approach can produce solutions that meet the mission requirements.
NASA Astrophysics Data System (ADS)
Baturin, A. P.; Chuvashov, I. N.
2006-03-01
The simplified method of setting of initial coordinates and velocities vector in the orbit fitting problem has been considered. The method is appropriate for the objects observed in a short arc. This method consists in the determination of coordinates and velocities by two close angular observations. The missing distance to the observed object is setting approximately. It is more convenient to compare with traditional methods of preliminary orbit calculation because it need not solving of eight-power equation. The testing of the method for several asteroids and comets has indicated that the accuracy of the resulting initial vector is sufficient for the further least-square orbit improvement.
Orbit Determination of Spacecraft in Earth-Moon L1 and L2 Libration Point Orbits
NASA Technical Reports Server (NTRS)
Woodard, Mark; Cosgrove, Daniel; Morinelli, Patrick; Marchese, Jeff; Owens, Brandon; Folta, David
2011-01-01
The ARTEMIS mission, part of the THEMIS extended mission, is the first to fly spacecraft in the Earth-Moon Lissajous regions. In 2009, two of the five THEMIS spacecraft were redeployed from Earth-centered orbits to arrive in Earth-Moon Lissajous orbits in late 2010. Starting in August 2010, the ARTEMIS P1 spacecraft executed numerous stationkeeping maneuvers, initially maintaining a lunar L2 Lissajous orbit before transitioning into a lunar L1 orbit. The ARTEMIS P2 spacecraft entered a L1 Lissajous orbit in October 2010. In April 2011, both ARTEMIS spacecraft will suspend Lissajous stationkeeping and will be maneuvered into lunar orbits. The success of the ARTEMIS mission has allowed the science team to gather unprecedented magnetospheric measurements in the lunar Lissajous regions. In order to effectively perform lunar Lissajous stationkeeping maneuvers, the ARTEMIS operations team has provided orbit determination solutions with typical accuracies on the order of 0.1 km in position and 0.1 cm/s in velocity. The ARTEMIS team utilizes the Goddard Trajectory Determination System (GTDS), using a batch least squares method, to process range and Doppler tracking measurements from the NASA Deep Space Network (DSN), Berkeley Ground Station (BGS), Merritt Island (MILA) station, and United Space Network (USN). The team has also investigated processing of the same tracking data measurements using the Orbit Determination Tool Kit (ODTK) software, which uses an extended Kalman filter and recursive smoother to estimate the orbit. The orbit determination results from each of these methods will be presented and we will discuss the advantages and disadvantages associated with using each method in the lunar Lissajous regions. Orbit determination accuracy is dependent on both the quality and quantity of tracking measurements, fidelity of the orbit force models, and the estimation techniques used. Prior to Lissajous operations, the team determined the appropriate quantity of tracking measurements that would be needed to meet the required orbit determination accuracies. Analysts used the Orbit Determination Error Analysis System (ODEAS) to perform covariance analyses using various tracking data schedules. From this analysis, it was determined that 3.5 hours of DSN TRK-2-34 range and Doppler tracking data every other day would suffice to meet the predictive orbit knowledge accuracies in the Lissajous region. The results of this analysis are presented. Both GTDS and ODTK have high-fidelity environmental orbit force models that allow for very accurate orbit estimation in the lunar Lissajous regime. These models include solar radiation pressure, Earth and Moon gravity models, third body gravitational effects from the Sun, and to a lesser extent third body gravitational effects from Jupiter, Venus, Saturn, and Mars. Increased position and velocity uncertainties following each maneuver, due to small execution performance errors, requires that several days of post-maneuver tracking data be processed to converge on an accurate post-maneuver orbit solution. The effects of maneuvers on orbit determination accuracy will be presented, including a comparison of the batch least squares technique to the extended Kalman filter/smoother technique. We will present the maneuver calibration results derived from processing post-maneuver tracking data. A dominant error in the orbit estimation process is the uncertainty in solar radiation pressure and the resultant force on the spacecraft. An estimation of this value can include many related factors, such as the uncertainty in spacecraft reflectivity and surface area which is a function of spacecraft orientation (spin-axis attitude), uncertainty in spacecraft wet mass, and potential seasonal variability due to the changing direction of the Sun line relative to the Earth-Moon Lissajous reference frame. In addition, each spacecraft occasionally enters into Earth or Moon penumbra or umbra and these shadow crossings reduche solar radiation force for several hours. The effects of these events on orbit determination accuracy will be presented. In order to plan for upcoming stationkeeping maneuvers, the maneuver planning team must take the current orbit estimate, propagate it forward to the planned maneuver time, and determine the optimal maneuver to maintain the Lissajous orbit for one or more revolutions. The propagation is performed using a Runge-Kutta 7/8 integrator and typically the position and velocity uncertainty increases with propagation time, increasing the overall uncertainty of the orbit state at the maneuver execution time. The effect of orbit knowledge uncertainty on stationkeeping operations will be presented.
Hill equations for satellite orbit determination
NASA Astrophysics Data System (ADS)
Vancoevorden, R. G.
1992-11-01
Equations of motion using a periodical circular orbit are addressed. The orbital perturbations are given with respect to this moving triad. This set of equations, called the Hill equations, exists of three second order linear differential equations. They describe the problem in a first order approximation. Depending on the type of the disturbing forces, there exist different solutions of these equations. When there are no disturbing forces, the equations are called the homogeneous Hill equations, and only the initial values of the state vector can change the shape of the orbit. Disturbing forces which are more complex, can be transformed into Fourier series and then used in the equations to get an exact analytical solution of the approximated problem. By looking at the complete solution of the Hill equations, it can be seen that there are a few cases in which the solutions are not valid. The so called critical frequencies give a resonant effect on the orbital perturbations. These frequencies are the zero frequency and the once per revolution frequency. Resonant sources are for instance: drag, solar pressure, etc. A simple rendezvous problem, which describes the use of the homogeneous equations of motion, is discussed, some resonant sources are explained, and two examples of some relativistic effects are given. The cases in which the disturbing frequency is almost equal to a critical frequency are described. The amplitudes of the perturbations can grow very big in these so called near resonance cases. As a result of this work, the Hill equations can be said to be very good for educative purposes, because they give a very good view on the effects of disturbing forces on the orbit of a satellite. It should always be kept in mind that many simplifications are made when deriving the Hill equations.
Tethered body problems and relative motion orbit determination
NASA Technical Reports Server (NTRS)
Eades, J. B., Jr.; Wolf, H.
1972-01-01
Selected problems dealing with orbiting tethered body systems have been studied. In addition, a relative motion orbit determination program was developed. Results from these tasks are described and discussed. The expected tethered body motions were examined, analytically, to ascertain what influence would be played by the physical parameters of the tether, the gravity gradient and orbit eccentricity. After separating the motion modes these influences were determined; and, subsequently, the effects of oscillations and/or rotations, on tether force, were described. A study was undertaken, by examining tether motions, to see what type of control actions would be needed to accurately place a mass particle at a prescribed position relative to a main vehicle. Other applications for tethers were studied. Principally these were concerned with the producing of low-level gee forces by means of stabilized tether configurations; and, the initiation of free transfer trajectories from tether supported vehicle relative positions.
Formation Flying In Highly Elliptical Orbits Initializing the Formation
NASA Technical Reports Server (NTRS)
Mailhe, Laurie; Schiff, Conrad; Hughes, Steven
2000-01-01
In this paper several methods are examined for initializing formations in which all spacecraft start in a common elliptical orbit subsequent to separation from the launch vehicle. The tetrahedron formation used on missions such as the Magnetospheric Multiscale (MMS), Auroral Multiscale Midex (AMM), and Cluster is used as a test bed Such a formation provides full three degrees-of-freedom in the relative motion about the reference orbit and is germane to several missions. The type of maneuver strategy that can be employed depends on the specific initial conditions of each member of the formation. Single-impulse maneuvers based on a Gaussian variation-of-parameters (VOP) approach, while operationally simple and intuitively-based, work only in a limited sense for a special class of initial conditions. These 'tailored' initial conditions are characterized as having only a few of the Keplerian elements different from the reference orbit. Attempts to achieve more generic initial conditions exceed the capabilities of the single impulse VOP. For these cases, multiple-impulse implementations are always possible but are generally less intuitive than the single-impulse case. The four-impulse VOP formalism discussed by Schaub is examined but smaller delta-V costs are achieved in our test problem by optimizing a Lambert solution.
Using Onboard Telemetry for MAVEN Orbit Determination
NASA Technical Reports Server (NTRS)
Lam, Try; Trawny, Nikolas; Lee, Clifford
2013-01-01
Determination of the spacecraft state has been traditional done using radiometric tracking data before and after the atmosphere drag pass. This paper describes our approach and results to include onboard telemetry measurements in addition to radiometric observables to refine the reconstructed trajectory estimate for the Mars Atmosphere and Volatile Evolution Mission (MAVEN). Uncertainties in the Mars atmosphere models, combined with non-continuous tracking degrade navigation accuracy, making MAVEN a key candidate for using onboard telemetry data to help complement its orbit determination process.
Orbit determination of Tance-1 satellite using VLBI data
NASA Astrophysics Data System (ADS)
Huang, Y.; Hu, X. G.; Huang, C.; Jiang, D. R.
2006-01-01
On 30 December, 2003, China successfully launched the first satellite Tance-1 of Chinese Geospace Double Star Exploration Program, i.e. "Double Star Program (DSP)", on an improved Long March 2C launch vehicle. The Tance-1 satellite is operating at an orbit around the earth with a 550km perigee, 78000km apogee and 28.5 degree inclination.VLBI technique can track Tance-1 satellite or even far satellites such as lunar vehicles. To validate the VLBI technique in the on-going Chinese lunar exploration mission, Shanghai Astronomical Observatory (SHAO) organized to track the Tance-1 satellite with Chinese three VLBI stations: Shanghai, Kunming and Urumchi Orbit Determination (OD) of the Tance-1 satellite with about two days VLBI dada, and the capability of OD with VLBI data are studied. The results show that the VLBI-based orbit solutions improve the fit level over the initial orbit. The VLBI-delay-based orbit solution shows that the RMS of residuals of VLBI delay data is about 5.5m, and about 2.0cm/s for the withheld VLBI delay rate data. The VLBI-delay-rate-based orbit solution shows that the RMS of residuals of VLBI delay rate data is about 1.3cm/s, and about 29m for the withheld VLBI delay data. In the situation of orbit determination with VLBI delay and delay rate data with data sigma 5.5m and 1.3cm/s respectively, the RMS of residuals are 5.5,m and 2.0cm/s respectively. The simulation data assess the performance of the solutions. Considering the dynamic model errors of the Tance-1 satellite, the accuracy of the position is about km magnitude, and the accuracy of the velocity is about cm/s magnitude. The simulation work also show the dramatic accuracy improvement of OD with VLBI and USB combined.
From Ancient Paradoxes to Modern Orbit Determination
NASA Astrophysics Data System (ADS)
Giorgini, Jon D.
2008-09-01
In the 5th century BC, Zeno advanced a set of paradoxes to show motion and time are impossible, hence an illusion. The problem of motion has since driven much scientific thought and discovery, extending to Einstein's insights and the quantum revolution. To determine and predict the motion of remote objects within the solar system, a methodology has been refined over centuries. It integrates ideas from astronomy, physics, mathematics, measurement, and probability theory, having motivated most of those developments. Recently generalized and made numerically efficient, statistical orbit determination has made it possible to remotely fly Magellan and other spacecraft through the turbulent atmospheres of Venus and other planets while estimating atmospheric structure and internal mass distributions of the planet. Over limited time-scales, the methodology can predict the position of the Moon within a meter and asteroids within tens of meters -- their velocities at the millimeter per second level -- while characterizing the probable correctness of the prediction. Current software and networks disseminate such ephemeris information in moments; over the last 12 years, 10 million ephemerides have been provided by the Horizons system, at the request of 300000 different users. Applications range from ground and space telescope pointing to correlation with observations recorded on Babylonian cuneiform tablets. Rapid orbit updates are particularly important for planetary radars integrating weak small-body echoes moving quickly through the frequency spectrum due to relative motion. A loop is established in which the predicted delay-Doppler measurement and uncertainties are used to configure the radar. Both predictions are then compared to actual results, the asteroid or comet orbit solution improved, and the radar system optimally adjusted. Still, after 2500 years and tremendous descriptive success, there remain substantial problems understanding and predicting motion.
Determination of the orbits of inner Jupiter satellites
NASA Astrophysics Data System (ADS)
Avdyushev, V. A.; Ban'shikova, M. A.
2008-08-01
Some problems in determining the orbits of inner satellites associated with the complex behavior of the target function, which is strongly ravine and which possesses multiple minima in the case of the satellite orbit is determined based on fragmentary observations distributed over a rather long time interval, are studied. These peculiarities of the inverse problems are considered by the example of the dynamics of the inner Jupiter satellites: Amalthea, Thebe, Adrastea, and Metis. Numerical models of the satellite motions whose parameters were determined based on ground-based observations available at the moment to date have been constructed. A composite approach has been proposed for the effective search for minima of the target function. The approach allows one to obtain the respective evaluations of the orbital parameters only for several tens of iterations even in the case of very rough initial approximations. If two groups of observations are available (Adrastea), a formal minimization of the target function is shown to give a solution set, which is the best solution from the point of view of representation of the orbital motion, which is impossible to choose. Other estimates are given characterizing the specific nature of the inverse problems.
Real-time Sub-cm Differential Orbit Determination of two Low-Earth Orbiters with GPS Bias Fixing
NASA Technical Reports Server (NTRS)
Wu, Sien-Chong; Bar-Sever, Yoaz E.
2006-01-01
An effective technique for real-time differential orbit determination with GPS bias fixing is formulated. With this technique, only real-time GPS orbits and clocks are needed (available from the NASA Global Differential GPS System with 10-20 cm accuracy). The onboard, realtime orbital states of user satellites (few meters in accuracy) are used for orbit initialization and integration. An extended Kalman filter is constructed for the estimation of the differential orbit between the two satellites as well as a reference orbit, together with their associating dynamics parameters. Due to close proximity of the two satellites and of similar body shapes, the differential dynamics are highly common and can be tightly constrained which, in turn, strengthens the orbit estimation. Without explicit differencing of GPS data, double-differenced phase biases are formed by a transformation matrix. Integer-valued fixing of these biases are then performed which greatly strengthens the orbit estimation. A 9-day demonstration between GRACE orbits with baselines of approx.200 km indicates that approx.80% of the double-differenced phase biases can successfully be fixed and the differential orbit can be determined to approx.7 mm as compared to the results of onboard K-band ranging.
Bayesian statistical approach to binary asteroid orbit determination
NASA Astrophysics Data System (ADS)
Kovalenko, I.; Stoica, R. S.; Hestroffer, D.; Doressoundiram, A.
2015-10-01
Orbit determination from observations is one of the classical problems in celestial mechanics. Here we present a statistical approach to banary asteroids orbit determination based on the algorithm of Monte Carlo Markov Chain (MCMC). Furthermore, the present method can be used on the orbit determination in the Gaia mission program for the observations of binary asteroids.
Mars exploration rovers orbit determination system modeling
NASA Astrophysics Data System (ADS)
Wawrzyniak, Geoffrey; Baird, Darren; Graat, Eric; McElrath, Tim; Portock, Brian; Watkins, Michael
2006-06-01
From June 2003 to January 2004, two spinning spacecraft journeyed from Earth to Mars. A team of navigators at the Jet Propulsion Laboratory (JPL) accurately determined the orbits of both Mars Exploration Rovers, Spirit and Opportunity. For the navigation process to be successful, the team needed to know how nongravitational effects and how measurement system properties affected the trajectory and data modeling. To accomplish this, in addition to the standard gravitational and radiometric modeling of the spacecraft, a calibration was performed on each spacecraft to determine the amount of ?V that might occur during a turn, a high-fidelity solar-radiation-pressure model was created, the spin signature was removed from the tracking data, the station locations of the Deep Space Network were resurveyed, and a model of interplanetary charged particles was developed. The result of this effort was near-perfect accuracy, surpassing the tight atmospheric-entry requirements for navigation of both spacecraft.
OrbView-3 Initial On-Orbit Characterization
NASA Technical Reports Server (NTRS)
Ross, Kent; Blonski, Slawomir; Holekamp, Kara; Pagnutti, Mary; Zanoni, Vicki; Carver, David; Fendley, Debbie; Smith, Charles
2004-01-01
NASA at Stennis Space Center (SSC) established a Space Act Agreement with Orbital Sciences Corporation (OSC) and ORBIMAGE Inc. to collaborate on the characterization of the OrbView-3 system and its imagery products and to develop characterization techniques further. In accordance with the agreement, NASA performed an independent radiometric, spatial, and geopositional accuracy assessment of OrbView-3 imagery acquired before completion of the system's initial on-orbit checkout. OSC acquired OrbView-3 imagery over SSC from July 2003 through January 2004, and NASA collected ground reference information coincident with many of these acquisitions. After evaluating all acquisitions, NASA deemed two multispectral images and five panchromatic images useful for characterization. NASA then performed radiometric, spatial, and geopositional characterizations.
Wincs/Swats Initial on-Orbit Performance Results
NASA Astrophysics Data System (ADS)
Nicholas, A. C.; Herrero, F. A.; Stephan, A. W.; Finne, T. T.
2014-12-01
The Winds-Ions-Neutral Composition Suite (WINCS) instrument, also know as the Small Wind and Temperature Spectrometer (SWATS), was designed and developed jointly by the Naval Research Laboratory (NRL) and NASA/Goddard Space Flight Center (GSFC) for ionosphere-thermosphere investigations in orbit between 120 and 550 km altitude. The WINCS design provides the following measurements in a single package with a low Size, Weight, and Power (SWaP): 7.6 x 7.6 x 7.1 cm outer dimensions, 0.75 kg total mass, and about 1.3 Watt total power: neutral winds, neutral temperature, neutral density, neutral composition, ion drifts, ion temperature, ion density and ion composition. The instrument is currently operating on the International Space Station (Sep. 2013) and on the STP-Sat3 spacecraft (Nov. 2013). Initial on-orbit results of the instrument will be presented.
NASA Astrophysics Data System (ADS)
Maier, A.; Baur, O.; Krauss, S.
2014-04-01
This contribution deals with Precise Orbit Determination of the Lunar Reconnaissance Orbiter, which is tracked with optical laser ranges in addition to radiometric Doppler range-rates and range observations. The optimum parameterization is assessed by overlap analysis tests that indicate the inner precision of the computed orbits. Information about the very long wavelengths of the lunar gravity field is inferred from the spacecraft positions. The NASA software packages GEODYN II and SOLVE were used for orbit determination and gravity field recovery [1].
20 CFR 320.5 - Initial determinations.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 20 Employees' Benefits 1 2011-04-01 2011-04-01 false Initial determinations. 320.5 Section 320.5... INITIAL DETERMINATIONS UNDER THE RAILROAD UNEMPLOYMENT INSURANCE ACT AND REVIEWS OF AND APPEALS FROM SUCH DETERMINATIONS § 320.5 Initial determinations. An initial determination shall be made with respect to each...
WINCS/SWATS Initial On-Orbit Performance Results
NASA Astrophysics Data System (ADS)
Nicholas, A. C.; Stephan, A. W.; Finne, T. T.; Herrero, F.
2013-12-01
The Winds-Ions-Neutral Composition Suite (WINCS) instrument, also know as the Small Wind and Temperature Spectrometer (SWATS), was designed and developed jointly by the Naval Research Laboratory (NRL) and NASA/Goddard Space Flight Center (GSFC) for ionosphere-thermosphere investigations in orbit between 120 and 550 km altitude. The WINCS design provides the following measurements in a single package with a low Size, Weight, and Power (SWaP): 7.6 x 7.6 x 7.1 cm outer dimensions, 0.75 kg total mass, and about 1.3 Watt total power: neutral winds, neutral temperature, neutral density, neutral composition, ion drifts, ion temperature, ion density and ion composition. Initial on-orbit results of the first flight of the instrument will be presented. The flight, scheduled for Aug 2013, is on the International Space Station as STP-H4 the instrument complement and will be in a 51.6° inclination circular orbit at 400 km altitude. The instrument will also be on the Space Environment Nano-Satellite Experiment (SENSE) and the STPSat-3 satellites, both expected to launch in the fall of 2013.
36 CFR 902.60 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-07-01
... 36 Parks, Forests, and Public Property 3 2010-07-01 2010-07-01 false Initial determination. 902.60... INFORMATION ACT Time Limitations § 902.60 Initial determination. (a) An initial determination whether or not... workdays in accordance with § 902.62. (b) Upon making initial determination, the Administrative...
42 CFR 405.803 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 42 Public Health 2 2010-10-01 2010-10-01 false Initial determination. 405.803 Section 405.803....803 Initial determination. (a) Carriers make initial determinations regarding claims for benefits under Medicare Part B. (b) An initial determination for purposes of this subpart includes...
Routine operational and high-precision orbit determination of envisat
NASA Astrophysics Data System (ADS)
Zandbergen, R.; Otten, M.; Righetti, P. L.; Kuijper, D.; Dow, J. M.
2003-04-01
ESA's Earth observation satellite Envisat was successfully launched on 1 March 2002 by an Ariane-5 launcher, and ESOC immediately took over the task of determining and predicting the orbit using S-band tracking data, and optimising the manoeuvre sequence to bring the spacecraft into an orbit accurately phased with ERS-2. On-board, Envisat carries, among others, a radar altimeter, a DORIS instrument and a laser retro-reflector array (SLR). Data from these instruments are being used at ESOC for high-precision orbit determination, for verification of the routine orbit determination and for cross-comparison with orbits computed on-board by the DORIS navigator and with those delivered with the Envisat products. This paper presents the first consolidated results obtained for Envisat routine and high-precision orbit determination. All orbit determination and control activities were performed with the software package NAPEOS, which was developed in-house.
GPS as an orbit determination subsystems
NASA Technical Reports Server (NTRS)
Fennessey, Richard; Roberts, Pat; Knight, Robin; Vanvolkinburg, Bart
1995-01-01
This paper evaluates the use of Global Positioning System (GPS) receivers as a primary source of tracking data for low-Earth orbit satellites. GPS data is an alternative to using range, azimuth, elevation, and range-rate (RAER) data from the Air Force Satellite Control Network antennas, the Space Ground Link System (SGLS). This evaluation is applicable to missions such as Skipper, a joint U.S. and Russian atmosphere research mission, that will rely on a GPS receiver as a primary tracking data source. The Detachment 2, Space and Missile Systems Center's Test Support Complex (TSC) conducted the evaluation based on receiver data from the Space Test Experiment Platform Mission O (STEP-O) and Advanced Photovoltaic and Electronics Experiments (APEX) satellites. The TSC performed orbit reconstruction and prediction on the STEP-0 and APEX vehicles using GPS receiver navigation solution data, SGLS RAER data, and SGLS anglesonly (azimuth and elevation) data. For the STEP-O case, the navigation solution based orbits proved to be more accurate than SGLS RAER based orbits. For the APEX case, navigation solution based orbits proved to be less accurate than SGLS RAER based orbits for orbit prediction, and results for orbit reconstruction were inconclusive due to the lack of a precise truth orbit. After evaluating several different GPS data processing methods, the TSC concluded that using GPS navigation solution data is a viable alternative to using SGLS RAER data.
49 CFR 7.31 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 49 Transportation 1 2010-10-01 2010-10-01 false Initial determinations. 7.31 Section 7.31... Initial determinations. An initial determination whether to release a record requested pursuant to subpart... determination. If the determination is to grant the request, the desired record will be made available...
Bayesian Statistical Approach To Binary Asteroid Orbit Determination
NASA Astrophysics Data System (ADS)
Dmitrievna Kovalenko, Irina; Stoica, Radu S.
2015-08-01
Orbit determination from observations is one of the classical problems in celestial mechanics. Deriving the trajectory of binary asteroid with high precision is much more complicate than the trajectory of simple asteroid. Here we present a method of orbit determination based on the algorithm of Monte Carlo Markov Chain (MCMC). This method can be used for the preliminary orbit determination with relatively small number of observations, or for adjustment of orbit previously determined.The problem consists on determination of a conditional a posteriori probability density with given observations. Applying the Bayesian statistics, the a posteriori probability density of the binary asteroid orbital parameters is proportional to the a priori and likelihood probability densities. The likelihood function is related to the noise probability density and can be calculated from O-C deviations (Observed minus Calculated positions). The optionally used a priori probability density takes into account information about the population of discovered asteroids. The a priori probability density is used to constrain the phase space of possible orbits.As a MCMC method the Metropolis-Hastings algorithm has been applied, adding a globally convergent coefficient. The sequence of possible orbits derives through the sampling of each orbital parameter and acceptance criteria.The method allows to determine the phase space of every possible orbit considering each parameter. It also can be used to derive one orbit with the biggest probability density of orbital elements.
Dynamic orbit determination using GPS measurements from TOPEX/POSEIDON
NASA Technical Reports Server (NTRS)
Schultz, B. E.; Tapley, B. D.; Abusali, P. A. M.; Rim, H. J.
1994-01-01
The GPAS data acquired by the TOPEX/POSEIDON (T/P) Demonstration Receiver (DR) have been used in a dynamic orbit determination, which was based on the description of the gravitational and nongravitational forces in the equations of motion. The GPS carrier phase data were processed in a double difference mode to remove clock errors, including the effects of Selective Availability. Simultaneous estimation of the T/P orbit and GPS orbits was performed using five 10-day cycles in the interval between December (1992) and April (1993). The resulting T/P orbits have been compared with the orbits determined from Satellite Laser Ranging, the French one-way Doppler tracking system, DORIS, and with the JPL reduced dynamic orbit determination strategies and force models with the GPS/DR to those used with SLR/DORIS, the radial component of the T/P orbit (based on JGM-2) was found to agree better than 30 mm (rms) and 35 mm with the JPL reduced dynamic orbit. An experiment gravity tuning was accomplished using four cycles of GPS/DR data. The resulting GPS./DR-orbits, determined by the dynamic technique with the experimental gravity field, are in better agreement with the JPL reduced dynamic orbits in both the radial component (21-25 mm) and altimeter crossover residuals than the JGM-2 orbits. (21-25 mm) and altimeter crossover residuals than the JGM-2 orbits.
Ulysses orbit determination at high declinations
NASA Technical Reports Server (NTRS)
Mcelrath, Timothy P.; Lewis, George D.
1995-01-01
The trajectory of the Ulysses spacecraft caused its geocentric declination to exceed 60 deg South for over two months during the Fall of 1994, permitting continuous tracking from a single site. During this time, spacecraft operations constraints allowed only Doppler tracking data to be collected, and imposed a high radial acceleration uncertainty on the orbit determination process. The unusual aspects of this situation have motivated a re-examination of the Hamilton-Melbourne results, which have been used before to estimate the information content of Doppler tracking for trajectories closer to the ecliptic. The addition of an acceleration term to this equation is found to significantly increase the declination uncertainty for symmetric passes. In addition, a simple means is described to transform the symmetric results when the tracking pass is non-symmetric. The analytical results are then compared against numerical studies of this tracking geometry and found to be in good agreement for the angular uncertainties. The results of this analysis are applicable to the Near Earth Asteroid Rendezvous (NEAR) mission and to any other missions with high declination trajectories, as well as to missions using short tracking passes and/or one-way Doppler data.
19 CFR 210.42 - Initial determinations.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 19 Customs Duties 3 2011-04-01 2011-04-01 false Initial determinations. 210.42 Section 210.42... TRADE ADJUDICATION AND ENFORCEMENT Determinations and Actions Taken § 210.42 Initial determinations. (a... administrative law judge shall certify the record to the Commission and shall file an initial determination...
Orbit determination and prediction study for Dynamic Explorer 2
NASA Technical Reports Server (NTRS)
Smith, R. L.; Nakai, Y.; Doll, C. E.
1983-01-01
Definitive orbit determination accuracy and orbit prediction accuracy for the Dynamic Explorer-2 (DE-2) are studied using the trajectory determination system for the period within six weeks of spacecraft reentry. Baseline accuracies using standard orbit determination models and methods are established. A promising general technique for improving the orbit determination accuracy of high drag orbits, estimation of random drag variations at perigee passages, is investigated. This technique improved the fit to the tracking data by a factor of five and improved the solution overlap consistency by a factor of two during a period in which the spacecraft perigee altitude was below 200 kilometers. The results of the DE-2 orbit predictions showed that improvement in short term prediction accuracy reduces to the problem of predicting future drag scale factors: the smoothness of the solar 10.7 centimeter flux density suggests that this may be feasible.
Two Line Element Aided Orbit Determination Using Single Station SLR Data
NASA Astrophysics Data System (ADS)
Liang, Z. P.; Liu, C. Z.; Fan, C. B.; Sun, M. G.
2012-03-01
It is difficult to use the single-station satellite laser ranging (SLR) data for orbit determination, due to the singular geometrical distribution of the observations. The single-station data generated by performing diffuse-reflection SLR to the orbital space debris are therefore ineffective for orbit improvement. We propose a method to resolve the singularity in the observation distribution. Since the initial orbits of space debris such as the two line elements (TLE) exist prior to the SLR tracking, we use it to simulate observations from other SLR sites. We combine the simulated and actual observations with a proper weight to fit an orbit, thus resolving the singularity in the observation distribution. We then propagate the fitted orbit forward in time to validate against the precision ephemeris derived from the international laser ranging service (ILRS). The method is implemented and applied to the satellite Ajisai. Using the single-station SLR data of five passes in one day and corresponding TLE as the initial orbit, we fit the orbit and the generated predictions. The predicted position error is less than 40 meter in five-day span, significantly improved over the initial SGP4 propagated orbit. The method's potential application to space debris orbit improvement is also discussed.
Phenomenological Determination of the Orbital Angular Momentum
Ramsey, Gordon P.
2009-08-04
Measurements involving the gluon spin, {delta}G(x, t) and the corresponding asymmetry, A(x,t) = {delta}G(x,t)/G(x,t) play an important role in quantitative understanding of proton structure. We have modeled the asymmetry perturbatively and calculated model corrections to obtain information about non-perturbative spin-orbit effects. These models are consistent with existing COMPASS and HERMES data on the gluon asymmetry. The J{sub z} = (1/2) sum rule is used to generate values of orbital angular momentum at LO and NLO. For models consistent with data, the orbital angular momentum is small. Our studies specify accuracy that future measurements should achieve to constrain theoretical models for nucleon structure.
20 CFR 725.420 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 3 2010-04-01 2010-04-01 false Initial determinations. 725.420 Section 725... Initial determinations. (a) Section 9501(d)(1)(A)(1) of the Internal Revenue Code (26 U.S.C.) provides... within 30 days after the date of an initial determination of eligibility by the Secretary. For...
19 CFR 207.114 - Initial determination.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 19 Customs Duties 3 2011-04-01 2011-04-01 false Initial determination. 207.114 Section 207.114... and Committee Proceedings § 207.114 Initial determination. (a) Time for filing of initial determination. (1) Except as may otherwise be ordered by the Commission, within ninety (90) days of the date...
Benefits Derived From Laser Ranging Measurements for Orbit Determination of the GPS Satellite Orbit
NASA Technical Reports Server (NTRS)
Welch, Bryan W.
2007-01-01
While navigation systems for the determination of the orbit of the Global Position System (GPS) have proven to be very effective, the current research is examining methods to lower the error in the GPS satellite ephemerides below their current level. Two GPS satellites that are currently in orbit carry retro-reflectors onboard. One notion to reduce the error in the satellite ephemerides is to utilize the retro-reflectors via laser ranging measurements taken from multiple Earth ground stations. Analysis has been performed to determine the level of reduction in the semi-major axis covariance of the GPS satellites, when laser ranging measurements are supplemented to the radiometric station keeping, which the satellites undergo. Six ground tracking systems are studied to estimate the performance of the satellite. The first system is the baseline current system approach which provides pseudo-range and integrated Doppler measurements from six ground stations. The remaining five ground tracking systems utilize all measurements from the current system and laser ranging measurements from the additional ground stations utilized within those systems. Station locations for the additional ground sites were taken from a listing of laser ranging ground stations from the International Laser Ranging Service. Results show reductions in state covariance estimates when utilizing laser ranging measurements to solve for the satellite s position component of the state vector. Results also show dependency on the number of ground stations providing laser ranging measurements, orientation of the satellite to the ground stations, and the initial covariance of the satellite's state vector.
Semi-Major Axis Knowledge and GPS Orbit Determination
NASA Technical Reports Server (NTRS)
Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)
2000-01-01
In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.
Semi-Major Axis Knowledge and GPS Orbit Determination
NASA Technical Reports Server (NTRS)
Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)
2000-01-01
In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning, Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.
Strategies for high-precision Global Positioning System orbit determination
NASA Technical Reports Server (NTRS)
Lichten, Stephen M.; Border, James S.
1987-01-01
Various strategies for the high-precision orbit determination of the GPS satellites are explored using data from the 1985 GPS field test. Several refinements to the orbit determination strategies were found to be crucial for achieving high levels of repeatability and accuracy. These include the fine tuning of the GPS solar radiation coefficients and the ground station zenith tropospheric delays. Multiday arcs of 3-6 days provided better orbits and baselines than the 8-hr arcs from single-day passes. Highest-quality orbits and baselines were obtained with combined carrier phase and pseudorange solutions.
Astrodynamics. Volume 1 - Orbit determination, space navigation, celestial mechanics.
NASA Technical Reports Server (NTRS)
Herrick, S.
1971-01-01
Essential navigational, physical, and mathematical problems of space exploration are covered. The introductory chapters dealing with conic sections, orientation, and the integration of the two-body problem are followed by an introduction to orbit determination and design. Systems of units and constants, as well as ephemerides, representations, reference systems, and data are then dealt with. A detailed attention is given to rendezvous problems and to differential processes in observational orbit correction, and in rendezvous or guidance correction. Finally, the Laplacian methods for determining preliminary orbits, and the orbit methods of Lagrange, Gauss, and Gibbs are reviewed.
Routine operational and high-precision orbit determination of Envisat
NASA Astrophysics Data System (ADS)
Zandbergen, R.; Righetti, P.; Otten, M.; Kuijper, D.; Dow, J.
ESA's Earth observation satellite Envisat was successfully launched on 1 March 2002 by an Ariane-5 launcher, and ESOC immediately took over the task of determining and predicting the orbit using S-band tracking data, and optimising the manoeuvre sequence to bring the spacecraft into an orbit accurately phased with ERS-2. On-board, Envisat carries, among others, a radar altimeter, a Doris instrument and an SLR retroreflector array. Data from these instruments are being used at ESOC for high -precision orbit determination, for verification of the routine orbit determination and for cross-comparison with orbits computed on-board by the Doris navigator and with those delivered with the Envisat products. This paper presents the first consolidated results obtained for Envisat routine and high - precision orbit determination. All orbit determination and control activities were performed with the software package Napeos, which was developed in-house. The future use of Napeos for orbit determination of satellites equipped with on -board GNSS receivers is also briefly addressed.
An orbit determination from debris impacts on measurement satellites
NASA Astrophysics Data System (ADS)
Fujita, Koki; Tasaki, Mitsuhiko; Furumoto, Masahiro; Hanada, Toshiya
2016-01-01
This work proposes a method to determine orbital plane of a micron-sized space debris cloud utilizing their impacts on measurement satellites. Given that debris impacts occur on a line of intersection between debris and satellites orbital planes, a couple of debris orbital parameters, right ascension of the ascending node, inclination, and nodal regression rate can be determined by impact times and locations measured from more than two satellites in different earth orbits. This paper proves that unique solution for the debris orbital parameters is obtained from the measurement data, and derives a computational scheme to estimate them. The effectiveness of the proposed scheme is finally demonstrated by a simulation test, in which measurement data are obtained from a numerical simulation considering realistic debris' and satellites' orbits.
ADEOS-II precise orbit determinations with GPS and SLR
NASA Astrophysics Data System (ADS)
Sawabe, M.; Kashimoto, M.
1999-01-01
In order to meet the scientific mission requirements for high accuracy orbit determination strongly required by future earth observation missions in the early 2000's, NASDA's precise orbit determination system with GPS (Global Positioning System) and SLR (Satellite Laser Ranging) is now being developed at the NASDA/TACC (Tracking and Control Center). This system, which is called GUTS (Global and high accUracy Trajectory determination System), will be used to demonstrate precise orbit determination with ADEOS-II (Advanced Earth Observing Satellite II), set for launch in the year 2000. This paper presents an overview of the GUTS and its experimental and operational plans.
NASA Astrophysics Data System (ADS)
LÃ¶cher, Anno; Kusche, JÃ¼rgen
2014-05-01
The Lunar Reconnaissance Orbiter (LRO) launched in 2009 by the National Aeronautics and Space Administration (NASA) still orbits the Moon in a polar orbit at an altitude of 50 kilometers and below. Its main objective is the detailed exploration of the Moon's surface by means of the Lunar Orbiter Laser Altimeter (LOLA) and three high resolution cameras bundled in the Lunar Reconnaissance Orbiter Camera (LROC) unit. Referring these observations to a Moon-fixed reference frame requires the computation of highly accurate and consistent orbits. For this task only Earth-based observations are available, primarily radiometric tracking data from stations in the United States, Australia and Europe. In addition, LRO is prepared for one-way laser measurements from specially adapted sites. Currently, 10 laser stations participate more or less regularly in this experiment. For operational reasons, the official LRO orbits from NASA only include radiometric data so far. In this presentation, we investigate the benefit of the laser ranging data by feeding both types of observations in an integrated orbit determination process. All computations are performed by an in-house software development based on a dynamical approach improving orbit and force parameters in an iterative way. Special attention is paid to the determination of bias parameters, in particular of timing biases between radio and laser stations and the drift and aging of the LRO spacecraft clock. The solutions from the combined data set will be compared to radio- and laser-only orbits as well as to the NASA orbits. Further results will show how recent gravity field models from the GRAIL mission can improve the accuracy of the LRO orbits.
NASA Astrophysics Data System (ADS)
Gan, Q. B.
2012-07-01
Autonomous satellite orbit determination is a key technique in autonomous satellite navigation. Many kinds of technologies have been proposed to realize the autonomous satellite navigation, such as the star sensor, the Earth magnetometer, the occultation time survey, and the phase measurement of X-ray pulsar signals. This dissertation studies a method of autonomous satellite orbit determination using star sensor. Moreover, the method is extended to the autonomous navigation of satellite constellation and the space-based surveillance. In chapters 1 and 2, some usual time and reference systems are introduced. Then the principles of several typical autonomous navigation methods, and their merits and shortcomings are analyzed. In chapter 3, the autonomous satellite orbit determination using star sensor and infrared Earth sensor (IRES) is specifically studied, which is based on the status movement simulation, the stellar background observation from star sensor, and the Earth center direction survey from IRES. By simulating the low Earth orbit satellites and pseudo Geostationary Earth orbit (PGEO) satellites, the precision of position and speed with autonomous orbit determination using star sensor is obtained. Besides, the autonomous orbit determination using star sensor with double detectors is studied. According to the observation equation's characters, an optimized type of star sensor and IRES initial assembly model is proposed. In the study of the PGEO autonomous orbit determination, an efficient sampling frequency of measurements is promoted. The simulation results confirm that the autonomous satellite orbit determination using star sensor is feasible for satellites with all kinds of altitudes. In chapter 4, the method of autonomous satellite orbit determination using star sensor is extended to the autonomous navigation of mini-satellite constellation. Combining with the high-accuracy inter satellite links data, the precision of the determined orbit and constellation configuration is higher than that ever expected. In chapter 5, two related pre-project researches are developed with respect to the space-based satellite surveillance. One solves the un-convergence question in the preliminary orbit determination and finds an advantageous preliminary orbit determination using inter satellite angle measurement. In the other pre-project research, a creative space-based satellite surveillance model is proposed, which is based on the autonomous surveillance platform navigation. Using the star sensor's navigation data associated with the inter satellite angle measurement, the orbit parameters of the tracking space objects and the surveillance platform are determined. Compared to the available experiment results overseas, the preliminary orbit determination method and the autonomous navigation surveillance platform model are found to be feasible. The research will significantly contribute to the new conception of ``space awareness'', as well as our country's space security construction.
The Importance of Semi-Major Axis Knowledge in the Determination of Near-Circular Orbits
NASA Technical Reports Server (NTRS)
Carpenter, J. Russell; Schiesser, Emil R.
1998-01-01
Modem orbit determination has mostly been accomplished using Cartesian coordinates. This usage has carried over in recent years to the use of GPS for satellite orbit determination. The unprecedented positioning accuracy of GPS has tended to focus attention more on the system's capability to locate the spacecraft's location at a particular epoch than on its accuracy in determination of the orbit, per se. As is well-known, the latter depends on a coordinated knowledge of position, velocity, and the correlation between their errors. Failure to determine a properly coordinated position/velocity state vector at a given epoch can lead to an epoch state that does not propagate well, and/or may not be usable for the execution of orbit adjustment maneuvers. For the quite common case of near-circular orbits, the degree to which position and velocity estimates are properly coordinated is largely captured by the error in semi-major axis (SMA) they jointly produce. Figure 1 depicts the relationships among radius error, speed error, and their correlation which exist for a typical low altitude Earth orbit. Two familiar consequences are the relationship Figure 1 shows are the following: (1) downrange position error grows at the per orbit rate of 3(pi) times the SMA error; (2) a velocity change imparted to the orbit will have an error of (pi) divided by the orbit period times the SMA error. A less familiar consequence occurs in the problem of initializing the covariance matrix for a sequential orbit determination filter. An initial covariance consistent with orbital dynamics should be used if the covariance is to propagate well. Properly accounting for the SMA error of the initial state in the construction of the initial covariance accomplishes half of this objective, by specifying the partition of the covariance corresponding to down-track position and radial velocity errors. The remainder of the in-plane covariance partition may be specified in terms of the flight path angle error of the initial state. Figure 2 illustrates the effect of properly and not properly initializing a covariance. This figure was produced by propagating the covariance shown on the plot, without process noise, in a circular low Earth orbit whose period is 5828.5 seconds. The upper subplot, in which the proper relationships among position, velocity, and their correlation has been used, shows overall error growth, in terms of the standard deviations of the inertial position coordinates, of about half of the lower subplot, whose initial covariance was based on other considerations.
Dependence of Orbit Determination Accuracy on the Observer Position
NASA Astrophysics Data System (ADS)
Vananti, Alessandro; Schildknecht, Thomas
2013-08-01
The Astronomical Institute of the University of Bern (AIUB) is conducting several search campaigns for space debris in Geostationary (GEO) and Medium Earth Orbits (MEO). Usually, to improve the quality of the determined orbits for newly discovered objects, follow-up observations are conducted. The latter take place at different times during the discovery night or in subsequent nights. The time interval between the observations plays an important role in the accuracy of the calculated orbits. Another essential parameter to consider is the position of the observer at the observation time. In this paper, the accuracy of the orbit determination with respect to the position of the observer is analyzed. The same observing site at varying epochs or multiple site locations involve different distances from the target object and a different observing angle with respect to its orbital plane and trajectory. The formal error in the orbit determination process is, among other dependencies, a function of the latter parameters. The analysis of this dependence is important to choose the appropriate observation strategy. One of the main questions that arises is e.g. whether observing the same object from different stations results in better determined orbits and, if yes, how big is the improvement. Another question is e.g. whether the observation from multiple sites needs to be simultaneous or not for a better orbit accuracy.
12 CFR 404.15 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 12 Banks and Banking 4 2010-01-01 2010-01-01 false Initial determination. 404.15 Section 404.15... the Privacy Act of 1974 § 404.15 Initial determination. (a) Time for processing. The Freedom of... the requester of any processing fee. (2) A denial is a determination to withhold any requested...
42 CFR 405.920 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 42 Public Health 2 2010-10-01 2010-10-01 false Initial determinations. 405.920 Section 405.920... PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED Determinations, Redeterminations, Reconsiderations, and Appeals Under Original Medicare (Part A and Part B) Initial Determinations § 405.920...
NASA Astrophysics Data System (ADS)
Janches, D.; Close, S.; Hormaechea, J. L.; Swarnalingam, N.; Murphy, A.; O'Connor, D.; Vandepeer, B.; Fuller, B.; Fritts, D. C.; Brunini, C.
2015-08-01
We present an initial survey in the southern sky of the sporadic meteoroid orbital environment obtained with the Southern Argentina Agile MEteor Radar (SAAMER) Orbital System (OS), in which over three-quarters of a million orbits of dust particles were determined from 2012 January through 2015 April. SAAMER-OS is located at the southernmost tip of Argentina and is currently the only operational radar with orbit determination capability providing continuous observations of the southern hemisphere. Distributions of the observed meteoroid speed, radiant, and heliocentric orbital parameters are presented, as well as those corrected by the observational biases associated with the SAAMER-OS operating parameters. The results are compared with those reported by three previous surveys performed with the Harvard Radio Meteor Project, the Advanced Meteor Orbit Radar, and the Canadian Meteor Orbit Radar, and they are in agreement with these previous studies. Weighted distributions for meteoroids above the thresholds for meteor trail electron line density, meteoroid mass, and meteoroid kinetic energy are also considered. Finally, the minimum line density and kinetic energy weighting factors are found to be very suitable for meteroid applications. The outcomes of this work show that, given SAAMERâ€™s location, the system is ideal for providing crucial data to continuously study the South Toroidal and South Apex sporadic meteoroid apparent sources.
Precise Orbit Determination of Low Earth Satellites at AIUB
NASA Astrophysics Data System (ADS)
Jaggi, A.; Bock, H.; Thaller, D.; Dach, R.; Beutler, G.; Prange, L.; Meyer, U.
2010-12-01
Many low Earth orbiting (LEO) satellites are nowadays equipped with on-board receivers to collect the observations from Global Navigation Satellite Systems (GNSS), such as the Global Positioning System (GPS), or with retro-reflectors for Satellite Laser Ranging (SLR). At the Astronomical Institute of the University of Bern (AIUB) LEO precise orbit determination (POD) using either GPS or SLR data is performed for satellites at very different altitudes. The classical numerical integration techniques used for dynamic orbit determination of LEO satellites at high altitudes are extended by pseudo-stochastic orbit modeling techniques for satellites at low altitudes to efficiently cope with force model deficiencies. Accuracies of a few centimeters are achieved by pseudo-stochastic orbit modeling, e.g., for the Gravity field and steady-state Ocean Circulation Explorer (GOCE).
The possible effect of reaction wheel unloading on orbit determination for Chang'E-1 lunar mission
NASA Astrophysics Data System (ADS)
Jianguo, Yan; Jingsong, Ping; Fei, Li
During the flight of 3-axis stabilized lunar orbiter i e SELENE main orbiter Chang E-1 due to the overflow of the accumulated angular momentum the reaction-wheel will be unloaded during certain period so as to release the angular momentum for initialization Then the momentum wheel will be reloaded for satellite attitude measurement and control Above action will not only change the attitude but also change the orbit of the spacecraft Assuming the reaction-wheel unloading is carried out twice a day according to the current engineering designation and plan for SELENE main orbiter and Chang E-1 missions considering the algebra configuration of the tracking stations the Moon and the lunar orbiter the orbit determination is simulated for 14 days evolution of lunar orbiter In the simulation the satellite orbit is generated using GEODYNII code Based on the generated orbit the common view time period of the satellite by VLBI and USB network in every day is computed the orbit determination is processed for all the arcs of the orbit The orbit determination result of 28 orbits in 14 days is provided The orbits cover most of the possible geometrical configuration among orbiter the Moon and the tracking network The analysis here can benefit the tracking designation and plan for Chang E-1 mission
Status of Precise Orbit Determination for Jason-2 Using GPS
NASA Technical Reports Server (NTRS)
Melachroinos, S.; Lemoine, F. G.; Zelensky, N. P.; Rowlands, D. D.; Pavlis, D. E.
2011-01-01
The JASON-2 satellite, launched in June 2008, is the latest follow-on to the successful TOPEX/Poseidon (T/P) and JASON-I altimetry missions. JASON-2 is equipped with a TRSR Blackjack GPS dual-frequency receiver, a laser retroreflector array, and a DORIS receiver for precise orbit determination (POD). The most recent time series of orbits computed at NASA GSFC, based on SLR/DORIS data have been completed using both ITRF2005 and ITRF2008. These orbits have been shown to agree radially at 1 cm RMS for dynamic vs SLRlDORIS reduced-dynamic orbits and in comparison with orbits produced by other analysis centers (Lemoine et al., 2010; Zelensky et al., 2010; Cerri et al., 2010). We have recently upgraded the GEODYN software to implement model improvements for GPS processing. We describe the implementation of IGS standards to the Jason2 GEODYN GPS processing, and other dynamical and measurement model improvements. Our GPS-only JASON-2 orbit accuracy is assessed using a number of tests including analysis of independent SLR and altimeter crossover residuals, orbit overlap differences, and direct comparison to orbits generated at GSFC using SLR and DORIS tracking, and to orbits generated externally at other centers. Tests based on SLR and the altimeter crossover residuals provide the best performance indicator for independent validation of the NASAlGSFC GPS-only reduced dynamic orbits. For the ITRF2005 and ITRF2008 implementation of our GPS-only obits we are using the IGS05 and IGS08 standards. Reduced dynamic versus dynamic orbit differences are used to characterize the remaining force model error and TRF instability. We evaluate the GPS vs SLR & DORIS orbits produced using the GEODYN software and assess in particular their consistency radially and the stability of the altimeter satellite reference frame in the Z direction for both ITRF2005 and ITRF2008 as a proxy to assess the consistency of the reference frame for altimeter satellite POD.
Real-time on-board orbit determination with DORIS
NASA Technical Reports Server (NTRS)
Berthias, J.-P.; Jayles, C.; Pradines, D.
1993-01-01
A spaceborne orbit determination system is being developed by the French Space Agency (CNES) for the SPOT 4 satellite. It processes DORIS measurements to produce an orbit with an accuracy of about 50O meters rms. In order to evaluate the reliability of the software, it was combined with the MERCATOR man/machine interface and used to process the TOPEX/Poseidon DORIS data in near real time during the validation phase of the instrument, at JPL and at CNES. This paper gives an overview of the orbit determination system and presents the results of the TOPEX/Poseidon experiment.
Evaluation of the IMP-16 microprocessor orbit determination system filter
NASA Technical Reports Server (NTRS)
Shenitz, C. M.; Tasaki, K. K.
1979-01-01
The results of the numerical tests performed in evaluating the interplanetary monitoring platform-16 orbit determination system are presented. The system is capable of performing orbit determination from satellite to satellite tracking data in applications technology satellite range and range rate format. The estimation scheme used is a Kalman filter, sequential (recursive) estimator. Descriptions of the tests performed and tabulations of the numerical results are included.
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2016-03-01
We present results for Precise Orbit Determination (POD) of the Lunar Reconnaissance Orbiter (LRO) based on two-way Doppler range-rates over a time span of ~13 months (January 3, 2011 to February 9, 2012). Different orbital arc lengths and various sets of empirical parameters were tested to seek optimal parametrization. An overlap analysis covering three months of Doppler data shows that the most precise orbits are obtained using an arc length of 2.5 days and estimating arc-wise constant empirical accelerations in along track direction. The overlap analysis over the entire investigated time span of 13 months indicates an orbital precision of 13.79 m, 14.17 m, and 1.28 m in along track, cross track, and radial direction, respectively, with 21.32 m in total position. We compare our orbits to the official science orbits released by the US National Aeronautics and Space Administration (NASA). The differences amount to 9.50 m, 6.98 m, and 1.50 m in along track, cross track, and radial direction, respectively, as well as 12.71 m in total position. Based on the reconstructed LRO orbits, we estimated lunar gravity field coefficients up to spherical harmonic degree and order 60. The results are compared to gravity field solutions derived from data collected by other lunar missions.
Space Capsule Recovery Orbit Determination System and Performance
NASA Astrophysics Data System (ADS)
Vighnesam, N. V.; Sonney, A.; Soni, P. K.
2008-08-01
Space Capsule Recovery (SRE), a small satellite, completely recoverable capsule was launched by the Polar Satellite Launch Vehicle (PSLV-C7) from the Indian spaceport Sriharikota on 10th January 2007 at 04:09UT along with Indian Remore Sensing Satellite CARTOSAT-2 and two micro satellites namely Nano- Peheunsat and Lapantubsat. The satellite was put into an almost nominal orbit of (630 X 638)km with an inclination of 97.94deg. The main objective of the SRE missions was to conduct microgravity experiment, de- orbit and recover it in Indian waters. The spacecraft was de-boosted after the payload operations in the micro- gravity environment. This was achieved in two steps. SRE was first placed from the injected circular orbit to Repetitive Elliptical Orbit (REO) and subsequently de- boosted for reentry and recovery. This paper describes the S-band based orbit determination system for SRE and its performance during different phases of the mission. Comparison of the inertial navigation system (INS) and nominal orbit with the achieved/estimated orbit was made. Orbit determination system was executed successfully through out the mission. Relatively large residues were observed in measurements during OD process due to continuous thruster activity through out the mission.
50 CFR 296.9 - Initial determination.
Code of Federal Regulations, 2011 CFR
2011-10-01
... 50 Wildlife and Fisheries 9 2011-10-01 2011-10-01 false Initial determination. 296.9 Section 296.9 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE CONTINENTAL SHELF FISHERMEN'S CONTINGENCY FUND § 296.9 Initial determination....
5 CFR 2502.7 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 5 Administrative Personnel 3 2010-01-01 2010-01-01 false Initial determination. 2502.7 Section 2502.7 Administrative Personnel OFFICE OF ADMINISTRATION, EXECUTIVE OFFICE OF THE PRESIDENT... § 2502.7 Initial determination. The General Counsel or his or her designee shall have the authority...
32 CFR 806.18 - Initial determinations.
Code of Federal Regulations, 2011 CFR
2011-07-01
... 32 National Defense 6 2011-07-01 2011-07-01 false Initial determinations. 806.18 Section 806.18 National Defense Department of Defense (Continued) DEPARTMENT OF THE AIR FORCE ADMINISTRATION AIR FORCE FREEDOM OF INFORMATION ACT PROGRAM § 806.18 Initial determinations. (a) Disclosure authorities make...
32 CFR 806.18 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-07-01
... 32 National Defense 6 2010-07-01 2010-07-01 false Initial determinations. 806.18 Section 806.18 National Defense Department of Defense (Continued) DEPARTMENT OF THE AIR FORCE ADMINISTRATION AIR FORCE FREEDOM OF INFORMATION ACT PROGRAM § 806.18 Initial determinations. (a) Disclosure authorities make...
28 CFR 301.305 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-07-01
... 28 Judicial Administration 2 2010-07-01 2010-07-01 false Initial determination. 301.305 Section... COMPENSATION Compensation for Work-Related Physical Impairment or Death § 301.305 Initial determination. A... delegated by the Board of Directors of Federal Prison Industries, Inc., pursuant to 28 CFR 0.99....
Precise orbit determination of the Lunar Reconnaissance Orbiter and first gravity field results
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2014-05-01
The Lunar Reconnaissance Orbiter (LRO) was launched in 2009 and is expected to orbit the Moon until the end of 2014. Among other instruments, LRO has a highly precise altimeter on board demanding an orbit accuracy of one meter in the radial component. Precise orbit determination (POD) is achieved with radiometric observations (Doppler range rates, ranges) on the one hand, and optical laser ranges on the other hand. LRO is the first satellite at a distance of approximately 360 000 to 400 000 km from the Earth that is routinely tracked with optical laser ranges. This measurement type was introduced to achieve orbits of higher precision than it would be possible with radiometric observations only. In this contribution we investigate the strength of each measurement type (radiometric range rates, radiometric ranges, optical laser ranges) based on single-technique orbit estimation. In a next step all measurement types are combined in a joined analysis. In addition to POD results, preliminary gravity field coefficients are presented being a subsequent product of the orbit determination process. POD and gravity field estimation was accomplished with the NASA/GSFC software packages GEODYN and SOLVE.
NASA Technical Reports Server (NTRS)
Lichten, S. M.; Estefan, J. A.
1990-01-01
Orbit covariance analyses pertaining to the Japanese VLBI Space Observatory Program (VSOP) MUSES-B satellite and to the International VLBI Satellite are presented. It is determined that a combination of Doppler and GPS measurements can provide the orbit accuracy required to support advanced radio interferometric experiments. For the VSOP, the required orbit accuracy of 130 m is easily met with two-way Doppler as the primary type of data; the 0.4 cm/s VSOP velocity requirement is also feasible provided that precise ground calibrations of tropospheric delays and station coordinates are available. It is concluded that combining the data from a VSOP GPS flight instrument with the ground GPS and two-way Doppler data will significantly enhance orbit determination accuracy in position and velocity.
Bayesian statistical approach to binary asteroid orbit determination
NASA Astrophysics Data System (ADS)
Kovalenko, Irina D.; Stoica, Radu S.; Emelyanov, N. V.; Doressoundiram, A.; Hestroffer, D.
2016-01-01
The problem of binary asteroids orbit determination is of particular interest, given knowledge of the orbit is the best way to derive the mass of the system. Orbit determination from observed points is a classic problem of celestial mechanics. However, in the case of binary asteroids, particularly with a small number of observations, the solution is not evident to derive. In the case of resolved binaries the problem consists in the determination of the relative orbit from observed relative positions of a secondary asteroid with respect to the primary. In this work, the problem is investigated as a statistical inverse problem. Within this context, we propose a method based on Bayesian modelling together with a global optimisation procedure that is based on the simulated annealing algorithm.
Robust Orbit Determination and Classification: A Learning Theoretic Approach
NASA Astrophysics Data System (ADS)
Sharma, S.; Cutler, J. W.
2015-11-01
Orbit determination involves estimation of a non-linear mapping from feature vectors associated with the position of the spacecraft to its orbital parameters. The de facto standard in orbit determination in real-world scenarios for spacecraft has been linearized estimators such as the extended Kalman filter. Such an estimator, while very accurate and convergent over its linear region, is hard to generalize over arbitrary gravitational potentials and diverse sets of measurements. It is also challenging to perform exact mathematical characterizations of the Kalman filter performance over such general systems. Here we present a new approach to orbit determination as a learning problem involving distribution regression and, also, for the multiple-spacecraft scenario, a transfer learning system for classification of feature vectors associated with spacecraft, and provide some associated analysis of such systems.
Precise orbit determination based on raw GPS measurements
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-Gürr, Torsten
2015-11-01
Precise orbit determination is an essential part of the most scientific satellite missions. Highly accurate knowledge of the satellite position is used to geolocate measurements of the onboard sensors. For applications in the field of gravity field research, the position itself can be used as observation. In this context, kinematic orbits of low earth orbiters (LEO) are widely used, because they do not include a priori information about the gravity field. The limiting factor for the achievable accuracy of the gravity field through LEO positions is the orbit accuracy. We make use of raw global positioning system (GPS) observations to estimate the kinematic satellite positions. The method is based on the principles of precise point positioning. Systematic influences are reduced by modeling and correcting for all known error sources. Remaining effects such as the ionospheric influence on the signal propagation are either unknown or not known to a sufficient level of accuracy. These effects are modeled as unknown parameters in the estimation process. The redundancy in the adjustment is reduced; however, an improvement in orbit accuracy leads to a better gravity field estimation. This paper describes our orbit determination approach and its mathematical background. Some examples of real data applications highlight the feasibility of the orbit determination method based on raw GPS measurements. Its suitability for gravity field estimation is presented in a second step.
Precise orbit determination based on raw GPS measurements
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-GÃ¼rr, Torsten
2016-03-01
Precise orbit determination is an essential part of the most scientific satellite missions. Highly accurate knowledge of the satellite position is used to geolocate measurements of the onboard sensors. For applications in the field of gravity field research, the position itself can be used as observation. In this context, kinematic orbits of low earth orbiters (LEO) are widely used, because they do not include a priori information about the gravity field. The limiting factor for the achievable accuracy of the gravity field through LEO positions is the orbit accuracy. We make use of raw global positioning system (GPS) observations to estimate the kinematic satellite positions. The method is based on the principles of precise point positioning. Systematic influences are reduced by modeling and correcting for all known error sources. Remaining effects such as the ionospheric influence on the signal propagation are either unknown or not known to a sufficient level of accuracy. These effects are modeled as unknown parameters in the estimation process. The redundancy in the adjustment is reduced; however, an improvement in orbit accuracy leads to a better gravity field estimation. This paper describes our orbit determination approach and its mathematical background. Some examples of real data applications highlight the feasibility of the orbit determination method based on raw GPS measurements. Its suitability for gravity field estimation is presented in a second step.
Distance-based relative orbital elements determination for formation flying system
NASA Astrophysics Data System (ADS)
He, Yanchao; Xu, Ming; Chen, Xi
2016-01-01
The present paper deals with determination of relative orbital elements based only on distance between satellites in the formation flying system, which has potential application in engineering, especially suited for rapid orbit determination required missions. A geometric simplification is performed to reduce the formation configuration in three-dimensional space to a plane. Then the equivalent actual configuration deviating from its nominal design is introduced to derive a group of autonomous linear equations on the mapping between the relative orbital elements differences and distance errors. The primary linear equations-based algorithm is initially proposed to conduct the rapid and precise determination of the relative orbital elements without the complex computation, which is further improved by least-squares method with more distance measurements taken into consideration. Numerical simulations and comparisons with traditional approaches are presented to validate the effectiveness of the proposed methods. To assess the performance of the two proposed algorithms, accuracy validation and Monte Carlo simulations are implemented in the presence of noises of distance measurements and the leader's absolute orbital elements. It is demonstrated that the relative orbital elements determination accuracy of two approaches reaches more than 90% and even close to the actual values for the least-squares improved one. The proposed approaches can be alternates for relative orbit determination without assistance of additional facilities in engineering for their fairly high efficiency with accuracy and autonomy.
Orbit Determination Accuracy for Comets on Earth-Impacting Trajectories
NASA Technical Reports Server (NTRS)
Kay-Bunnell, Linda
2004-01-01
The results presented show the level of orbit determination accuracy obtainable for long-period comets discovered approximately one year before collision with Earth. Preliminary orbits are determined from simulated observations using Gauss' method. Additional measurements are incorporated to improve the solution through the use of a Kalman filter, and include non-gravitational perturbations due to outgassing. Comparisons between observatories in several different circular heliocentric orbits show that observatories in orbits with radii less than 1 AU result in increased orbit determination accuracy for short tracking durations due to increased parallax per unit time. However, an observatory at 1 AU will perform similarly if the tracking duration is increased, and accuracy is significantly improved if additional observatories are positioned at the Sun-Earth Lagrange points L3, L4, or L5. A single observatory at 1 AU capable of both optical and range measurements yields the highest orbit determination accuracy in the shortest amount of time when compared to other systems of observatories.
UD filtering and smoothing applied to orbit determination
NASA Astrophysics Data System (ADS)
Kuga, Helio Koiti; Rios-Neto, Atair; Orlando, Valcir
1989-08-01
The performance is described of filtering and smoothing techniques applied to orbit determination of earth satellites. A (forward pass) Kalman filter along with an adaptive procedure for estimating the dynamic noise level which prevents the divergence of estimates due to inaccurate modelling of the orbital motion is implemented in the UD factorization form. The backward smoother is a numerically efficient version of the Rauch-Tung-Striebel (RTS) smoother. This smoother, developed in the UD form, has proven to be economical, compact and competitive for computer implementation. Digital simulations are performed for 2 situations of orbit determination: short arc orbit determination with favorably tracking geometry; and long arc orbit determination with existing tracking stations. Tests containing many levels of modelling degradation are carried out. The forward pass adaptive UD filter behaves so as to deal with these modelling nuisances and does not allow the divergence phenomenon to occur. In no real time operations, the backward UD smoother is used to improve the accuracy of the estimates and of the covariances resulting from the filtering phase. The results show, in the main, that the UD smoother can enhance the accuracy of the forward pass filter, sometimes by an order of magnitude. For post-flight analysis, the UD smoother is a useful tool when one aims at reconstituting the entire trajectory of the orbital motion covered by the tracking data.
Precise orbit determination and rapid orbit recovery supported by time synchronization
NASA Astrophysics Data System (ADS)
Guo, Rui; Zhou, JianHua; Hu, XiaoGong; Liu, Li; Tang, Bo; Li, XiaoJie; Wu, Shan
2015-06-01
In order to maintain optimal signal coverage, GNSS satellites have to experience orbital maneuvers. For China's COMPASS system, precise orbit determination (POD) as well as rapid orbit recovery after maneuvers contribute to the overall Positioning, Navigation and Timing (PNT) service performance in terms of accuracy and availability. However, strong statistical correlations between clock offsets and the radial component of a satellite's positions require long data arcs for POD to converge. We propose here a new strategy which relies on time synchronization between ground tracking stations and in-orbit satellites. By fixing satellite clock offsets measured by the satellite station two-way synchronization (SSTS) systems and receiver clock offsets, POD and orbital recovery performance can be improved significantly. Using the Satellite Laser Ranging (SLR) as orbital accuracy evaluation, we find the 4-hr recovered orbit achieves about 0.71 m residual root mean square (RMS) error of fit SLR data, the recovery time is improved from 24-hr to 4-hr compared with the conventional POD without time synchronization support. In addition, SLR evaluation shows that for 1-hr prediction, about 1.47 m accuracy is achieved with the new proposed POD strategy.
Evaluation of Improved Spacecraft Models for GLONASS Orbit Determination
NASA Astrophysics Data System (ADS)
Weiss, J. P.; Sibthorpe, A.; Harvey, N.; Bar-Sever, Y.; Kuang, D.
2010-12-01
High-fidelity spacecraft models become more important as orbit determination strategies achieve greater levels of precision and accuracy. In this presentation, we assess the impacts of new solar radiation pressure and attitude models on precise orbit determination (POD) for GLONASS spacecraft within JPLs GIPSY-OASIS software. A new solar radiation pressure model is developed by empirically fitting a Fourier expansion to solar pressure forces acting on the spacecraft X, Y, Z components using one year of recent orbit data. Compared to a basic “box-wing” solar pressure model, the median 24-hour orbit prediction accuracy for one month of independent test data improves by 43%. We additionally implement an updated yaw attitude model during eclipse periods. We evaluate the impacts of both models on post-processed POD solutions spanning 6-months. We consider a number of metrics such as internal orbit and clock overlaps as well as comparisons to independent solutions. Improved yaw attitude modeling reduces the dependence of these metrics on the “solar elevation” angle. The updated solar pressure model improves orbit overlap statistics by several mm in the median sense and centimeters in the max sense (1D). Orbit differences relative to the IGS combined solution are at or below the 5 cm level (1D RMS).
GRAIL Orbit Determination for the Science Phase and Extended Mission
NASA Technical Reports Server (NTRS)
Ryne, Mark; Antreasian, Peter; Broschart, Stephen; Criddle, Kevin; Gustafson, Eric; Jefferson, David; Lau, Eunice; Ying Wen, Hui; You, Tung-Han
2013-01-01
The Gravity Recovery and Interior Laboratory Mission (GRAIL) is the 11th mission of the NASA Discovery Program. Its objective is to help answer funda-mental questions about the Moon's internal structure, thermal evolution, and collisional history. GRAIL employs twin spacecraft, which fly in formation in low altitude polar orbits around the Moon. An improved global lunar gravity field is derived from high-precision range-rate measurements of the distance between the two spacecraft. The purpose of this paper is to describe the strategies used by the GRAIL Orbit Determination Team to overcome challenges posed during on-orbit operations.
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2015-04-01
The Lunar Reconnaissance Orbiter (LRO), launched in 2009, is well suited for the estimation of the long wavelengths of the lunar gravity field due to its low altitude of 50 km. Further, the orbit of LRO was polar for two years providing global coverage. The satellite has been primarily tracked via S-band (mainly two-way Doppler range-rates and two-way radiometric ranges) from the dedicated station in White Sands and from the Universal Space Network (USN). Due to the onboard altimeter the orbital precision requirement in the radial direction was rigorously defined as 1m. Because simulation studies before LRO's launch showed that this precision could not be reached with S-band observations alone, it was decided to additionally track LRO via optical laser ranges. It is worthwhile to point out that LRO is the first spacecraft in interplanetary space routinely tracked with optical one-way laser ranges. Gravity field recovery from orbit perturbations is intrinsically related to precise orbit determination. This is why considerable effort was made to find the optimum settings for orbit modeling. For a time span of three months we conducted a series of orbit overlapping tests based on Doppler observations to find the optimum arc length and the optimum set of empirical parameters. The analysis of observation residuals and orbit overlap differences showed that the estimated orbits are most precise when subdividing the time span into 2.5 days and estimating one constant empirical acceleration in along track direction. These settings were then used to analyze 13 months of Doppler data to LRO. The processing of the optical one-way laser was difficult due to the involvement of two non-synchronous clocks in one measurement (one clock at the ground station and one clock onboard LRO). The NASA software GEODYN, which was used for orbit determination and parameter estimation, models the LRO clock using a drift rate (first-order term) and an aging rate (second-order term). It seems, however, that this clock parametrization is not able to fully capture the signature posed on the measurement due to the two clocks. The precision of the orbits based solely on laser ranges is considerably lower compared to the Doppler-only orbits. For this reason, our lunar gravity field solution, which was estimated up to degree and order 60, is based solely on Doppler range-rates.
Intial orbit determination results for Jason-1: towards a 1-cm orbit
NASA Technical Reports Server (NTRS)
Haines, B. J.; Haines, B.; Bertiger, W.; Desai, S.; Kuang, D.; Munson, T.; Reichert, A.; Young, L.; Willis, P.
2002-01-01
The U.S/France Jason-1 oceanographic mission is carrying state-of-the-art radiometric tracking systems (GPS and Doris) to support precise orbit determination (POD) requirements. The performance of the systems is strongly reflected in the early POD results. Results of both internal and external (e.g., satellite laser ranging) comparisons support that the 2.5 cm radial Rh4S requirement is being readily met, and provide reasons for optimism that 1 cm can be achieved. We discuss the POD strategy underlying these orbits, as well as the challenging issues that bear on the understanding and characterization of an orbit solution at the l-cm level. We also describe a system for producing science quality orbits in near real time in order to support emerging applications in operational oceanography.
NASA Astrophysics Data System (ADS)
Ko, H.; Scheeres, D.
2014-09-01
Representing spacecraft orbit anomalies between two separate states is a challenging but an important problem in achieving space situational awareness for an active spacecraft. Incorporation of such a capability could play an essential role in analyzing satellite behaviors as well as trajectory estimation of the space object. A general way to deal with the anomaly problem is to add an estimated perturbing acceleration such as dynamic model compensation (DMC) into an orbit determination process based on pre- and post-anomaly tracking data. It is a time-consuming numerical process to find valid coefficients to compensate for unknown dynamics for the anomaly. Even if the orbit determination filter with DMC can crudely estimate an unknown acceleration, this approach does not consider any fundamental element of the unknown dynamics for a given anomaly. In this paper, a new way of representing a spacecraft anomaly using an interpolation technique with the Thrust-Fourier-Coefficients (TFCs) is introduced and several anomaly cases are studied using this interpolation method. It provides a very efficient way of reconstructing the fundamental elements of the dynamics for a given spacecraft anomaly. Any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver using an interpolation technique with the TFCs. Given unconnected orbit states between two epochs due to a spacecraft anomaly, it is possible to obtain a unique control law using the TFCs that is able to generate the desired secular behavior for the given orbital changes. This interpolation technique can capture the fundamental elements of combined unmodeled anomaly events. The interpolated orbit trajectory, using the TFCs compensating for a given anomaly, can be used to improve the quality of orbit fits through the anomaly period and therefore help to obtain a good orbit determination solution after the anomaly. Orbit Determination Toolbox (ODTBX) is modified to adapt this technique in order to verify the performance of this interpolation approach. Spacecraft anomaly cases are based on either single or multiple low or high thrust maneuvers and the unknown thrust accelerations are recovered and compared with the true thrust acceleration. The advantage of this approach is to easily append TFCs and its dynamics to the pre-built ODTBX, which enables us to blend post-anomaly tracking data to improve the performance of the interpolation representation in the absence of detailed information about a maneuver. It allows us to improve space situational awareness in the areas of uncertainty propagation, anomaly characterization and track correlation.
Orbit determination support of the Ocean Topography Experiment (TOPEX)/Poseidon operational orbit
NASA Technical Reports Server (NTRS)
Schanzle, A. F.; Rovnak, J. E.; Bolvin, D. T.; Doll, C. E.
1993-01-01
The Ocean Topography Experiment (TOPEX/Poseidon) mission is designed to determine the topography of the Earth's sea surface over a 3-year period, beginning shortly after launch in July 1992. TOPEX/Poseidon is a joint venture between the United States National Aeronautics and Space Administration (NASA) and the French Centre Nationale d'Etudes Spatiales. The Jet Propulsion Laboratory is NASA's TOPEX/Poseidon project center. The Tracking and Data Relay Satellite System (TDRSS) will nominally be used to support the day-to-day orbit determination aspects of the mission. Due to its extensive experience with TDRSS tracking data, the NASA Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will receive and process TDRSS observational data. To fulfill the scientific goals of the mission, it is necessary to achieve and maintain a very precise orbit. The most stringent accuracy requirements are associated with planning and evaluating orbit maneuvers, which will place the spacecraft in its mission orbit and maintain the required ground track. To determine if the FDF can meet the TOPEX/Poseidon maneuver accuracy requirements, covariance analysis was undertaken with the Orbit Determination Error Analysis System (ODEAS). The covariance analysis addressed many aspects of TOPEX/Poseidon orbit determination, including arc length, force models, and other processing options. The most recent analysis has focused on determining the size of the geopotential field necessary to meet the maneuver support requirements. Analysis was undertaken with the full 50 x 50 Goddard Earth Model (GEM) T3 field as well as smaller representations of this model.
Application of GPS tracking techniques to orbit determination for TDRS
NASA Technical Reports Server (NTRS)
Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S. C.
1993-01-01
In this paper, we evaluate two fundamentally different approaches to TDRS orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRSS spacecraft. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRSS spacecraft broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and TDRSS satellites from a small ground network. Both strategies can be designed to meet future operational requirements for TDRS-2 orbit determination.
Evaluation of semiempirical atmospheric density models for orbit determination applications
NASA Technical Reports Server (NTRS)
Cox, C. M.; Feiertag, R. J.; Oza, D. H.; Doll, C. E.
1994-01-01
This paper presents the results of an investigation of the orbit determination performance of the Jacchia-Roberts (JR), mass spectrometer incoherent scatter 1986 (MSIS-86), and drag temperature model (DTM) atmospheric density models. Evaluation of the models was performed to assess the modeling of the total atmospheric density. This study was made generic by using six spacecraft and selecting time periods of study representative of all portions of the 11-year cycle. Performance of the models was measured for multiple spacecraft, representing a selection of orbit geometries from near-equatorial to polar inclinations and altitudes from 400 kilometers to 900 kilometers. The orbit geometries represent typical low earth-orbiting spacecraft supported by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). The best available modeling and orbit determination techniques using the Goddard Trajectory Determination System (GTDS) were employed to minimize the effects of modeling errors. The latest geopotential model available during the analysis, the Goddard earth model-T3 (GEM-T3), was employed to minimize geopotential model error effects on the drag estimation. Improved-accuracy techniques identified for TOPEX/Poseidon orbit determination analysis were used to improve the Tracking and Data Relay Satellite System (TDRSS)-based orbit determination used for most of the spacecraft chosen for this analysis. This paper shows that during periods of relatively quiet solar flux and geomagnetic activity near the solar minimum, the choice of atmospheric density model used for orbit determination is relatively inconsequential. During typical solar flux conditions near the solar maximum, the differences between the JR, DTM, and MSIS-86 models begin to become apparent. Time periods of extreme solar activity, those in which the daily and 81-day mean solar flux are high and change rapidly, result in significant differences between the models. During periods of high geomagnetic activity, the standard JR model was outperformed by DTM. Modification of the JR model to use a geomagnetic heating delay of 3 hours, as used in DTM, instead of the 6.7-hour delay produced results comparable to or better than the DTM performance, reducing definitive orbit solution ephermeris overlap differences by 30 to 50 percent. The reduction in the overlap differences would be useful for mitigating the impact of geomagnetic storms on orbit prediction.
Precise orbit determination of Beidou Satellites at GFZ
NASA Astrophysics Data System (ADS)
Deng, Zhiguo; Ge, Maorong; Uhlemann, Maik; Zhao, Qile
2014-05-01
In December 2012 the Signal-In-Space Interface Control Document (ICD) of the BeiDou Navigation Satellite System (BeiDou system) was published. Currently the initial BeiDou regional navigation satellite system consisting of 14 satellites was completed, and provides observation data of five Geostationary-Earth-Orbit (GEO)satellites, five Inclined-GeoSynchronous-Orbit (IGSO) satellites and four Medium-Earth-Orbit (MEO) satellites. The Helmholtz Centre Potsdam GFZ German Research Centre for Geosciences (GFZ) contributes as one of the analysis centers to the International GNSS Service (IGS) since many years. In 2012 the IGS began the "Multi GNSS EXperiment" (MGEX), which supports the new GNSS, such as Galileo, Compass, and QZSS. Based on tracking data of BeiDou-capable receivers from the MGEX and chinese BeiDou networks up to 45 global distributed stations are selected to estimate orbit and clock parameters of the GPS/BeiDou satellites. Some selected results from the combined GPS/BeiDou data processing with 10 weeks of data from 2013 are shown. The quality of the orbit and clock products are assessed by means of orbit overlap statistics, clock stabilities as well as an independent validation with SLR measurements. At the end an outlook about GFZ AC's future Multi-GNSS activities will be given.
Precise Orbit Determination of BeiDou Navigation Satellite System
NASA Astrophysics Data System (ADS)
He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald
2013-04-01
China has been developing its own independent satellite navigation system since decades. Now the COMPASS system, also known as BeiDou, is emerging and gaining more and more interest and attention in the worldwide GNSS communities. The current regional BeiDou system is ready for its operational service around the end of 2012 with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit satellites (IGSO) and four Medium Earth orbit (MEO) satellites in operation. Besides the open service with positioning accuracy of around 10m which is free to civilian users, both precise relative positioning, and precise point positioning are demonstrated as well. In order to enhance the BeiDou precise positioning service, Precise Orbit Determination (POD) which is essential of any satellite navigation system has been investigated and studied thoroughly. To further improving the orbits of different types of satellites, we study the impact of network coverage on POD data products by comparing results from tracking networks over the Chinese territory, Asian-Pacific, Asian and of global scale. Furthermore, we concentrate on the improvement of involving MEOs on the orbit quality of GEOs and IGSOs. POD with and without MEOs are undertaken and results are analyzed. Finally, integer ambiguity resolution which brings highly improvement on orbits and positions with GPS data is also carried out and its effect on POD data products is assessed and discussed in detail. Seven weeks of BeiDou data from a ground tracking network, deployed by Wuhan University is employed in this study. The test constellation includes four GEO, five IGSO and two MEO satellites in operation. The three-day solution approach is employed to enhance its strength due to the limited coverage of the tracking network and the small movement of most of the satellites. A number of tracking scenarios and processing schemas are identified and processed and overlapping orbit differences are utilized to qualify the estimated orbits and clocks. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. For the current tracking network, deploying tracking stations on the eastern side, for example in New Zealand and/or in Hawaii, will significantly reduce along-track biases of GEOs on the same side. The involvement of MEOs and ambiguity-fixing also make the orbits better but rather moderate. Key words: BeiDou, precise orbit determination (POD), tracking network, ambiguity-fixing
Implementation of a low-cost, commercial orbit determination system
NASA Technical Reports Server (NTRS)
Corrigan, Jim
1994-01-01
Traditional satellite and launch control systems have consisted of custom solutions requiring significant development and maintenance costs. These systems have typically been designed to support specific program requirements and are expensive to modify and augment after delivery. The expanding role of space in today's marketplace combined with the increased sophistication and capabilities of modern satellites has created a need for more efficient, lower cost solutions to complete command and control systems. Recent technical advances have resulted in commercial-off-the-shelf products which greatly reduce the complete life-cycle costs associated with satellite launch and control system procurements. System integrators and spacecraft operators have, however, been slow to integrate these commercial based solutions into a comprehensive command and control system. This is due, in part, to a resistance to change and the fact that many available products are unable to effectively communicate with other commercial products. The United States Air Force, responsible for the health and safety of over 84 satellites via its Air Force Satellite Control Network (AFSCN), has embarked on an initiative to prove that commercial products can be used effectively to form a comprehensive command and control system. The initial version of this system is being installed at the Air Force's Center for Research Support (CERES) located at the National Test Facility in Colorado Springs, Colorado. The first stage of this initiative involved the identification of commercial products capable of satisfying each functional element of a command and control system. A significant requirement in this product selection criteria was flexibility and ability to integrate with other available commercial products. This paper discusses the functions and capabilities of the product selected to provide orbit determination functions for this comprehensive command and control system.
Implementation of a low-cost, commercial orbit determination system
NASA Astrophysics Data System (ADS)
Corrigan, Jim
1994-11-01
Traditional satellite and launch control systems have consisted of custom solutions requiring significant development and maintenance costs. These systems have typically been designed to support specific program requirements and are expensive to modify and augment after delivery. The expanding role of space in today's marketplace combined with the increased sophistication and capabilities of modern satellites has created a need for more efficient, lower cost solutions to complete command and control systems. Recent technical advances have resulted in commercial-off-the-shelf products which greatly reduce the complete life-cycle costs associated with satellite launch and control system procurements. System integrators and spacecraft operators have, however, been slow to integrate these commercial based solutions into a comprehensive command and control system. This is due, in part, to a resistance to change and the fact that many available products are unable to effectively communicate with other commercial products. The United States Air Force, responsible for the health and safety of over 84 satellites via its Air Force Satellite Control Network (AFSCN), has embarked on an initiative to prove that commercial products can be used effectively to form a comprehensive command and control system. The initial version of this system is being installed at the Air Force's Center for Research Support (CERES) located at the National Test Facility in Colorado Springs, Colorado. The first stage of this initiative involved the identification of commercial products capable of satisfying each functional element of a command and control system. A significant requirement in this product selection criteria was flexibility and ability to integrate with other available commercial products. This paper discusses the functions and capabilities of the product selected to provide orbit determination functions for this comprehensive command and control system.
Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation
NASA Astrophysics Data System (ADS)
Miller, J. K.; Antreasian, P. J.; Georgini, J.; Owen, W. M.; Williams, B. G.; Yeomans, D. K.
1995-01-01
Navigation of the orbit phase of the Near Earth steroid Rendezvous (NEAR) mission will re,quire determination of certain physical parameters describing the size, shape, gravity field, attitude and inertial properties of Eros. Prior to launch, little was known about Eros except for its orbit which could be determined with high precision from ground based telescope observations. Radar bounce and light curve data provided a rough estimate of Eros shape and a fairly good estimate of the pole, prime meridian and spin rate. However, the determination of the NEAR spacecraft orbit requires a high precision model of Eros's physical parameters and the ground based data provides only marginal a priori information. Eros is the principal source of perturbations of the spacecraft's trajectory and the principal source of data for determining the orbit. The initial orbit determination strategy is therefore concerned with developing a precise model of Eros. The original plan for Eros orbital operations was to execute a series of rendezvous burns beginning on December 20,1998 and insert into a close Eros orbit in January 1999. As a result of an unplanned termination of the rendezvous burn on December 20, 1998, the NEAR spacecraft continued on its high velocity approach trajectory and passed within 3900 km of Eros on December 23, 1998. The planned rendezvous burn was delayed until January 3, 1999 which resulted in the spacecraft being placed on a trajectory that slowly returns to Eros with a subsequent delay of close Eros orbital operations until February 2001. The flyby of Eros provided a brief glimpse and allowed for a crude estimate of the pole, prime meridian and mass of Eros. More importantly for navigation, orbit determination software was executed in the landmark tracking mode to determine the spacecraft orbit and a preliminary shape and landmark data base has been obtained. The flyby also provided an opportunity to test orbit determination operational procedures that will be used in February of 2001. The initial attitude and spin rate of Eros, as well as estimates of reference landmark locations, are obtained from images of the asteroid. These initial estimates are used as a priori values for a more precise refinement of these parameters by the orbit determination software which combines optical measurements with Doppler tracking data to obtain solutions for the required parameters. As the spacecraft is maneuvered; closer to the asteroid, estimates of spacecraft state, asteroid attitude, solar pressure, landmark locations and Eros physical parameters including mass, moments of inertia and gravity harmonics are determined with increasing precision. The determination of the elements of the inertia tensor of the asteroid is critical to spacecraft orbit determination and prediction of the asteroid attitude. The moments of inertia about the principal axes are also of scientific interest since they provide some insight into the internal mass distribution. Determination of the principal axes moments of inertia will depend on observing free precession in the asteroid's attitude dynamics. Gravity harmonics are in themselves of interest to science. When compared with the asteroid shape, some insight may be obtained into Eros' internal structure. The location of the center of mass derived from the first degree harmonic coefficients give a direct indication of overall mass distribution. The second degree harmonic coefficients relate to the radial distribution of mass. Higher degree harmonics may be compared with surface features to gain additional insight into mass distribution. In this paper, estimates of Eros physical parameters obtained from the December 23,1998 flyby will be presented. This new knowledge will be applied to simplification of Eros orbital operations in February of 2001. The resulting revision to the orbit determination strategy will also be discussed.
Evaluation of orbit determination using dual-TDRS tracking
NASA Technical Reports Server (NTRS)
Oza, D. H.; Hodjatzadeh, M.; Radomski, M. S.; Doll, C. E.; Gramling, C. J.
1990-01-01
This paper describes the results of a study to evaluate the orbit determinatioin of Tracking and Data Relay Satellite System (TDRSS) user spacecraft within the dual-Tracking and Data Relay Satellite (TDRS) environment. Dense TDRSS tracking of the Earth Radiation Budget Satellite (ERBS) was acquired for the period August 16 through 22, 1989. This tracking information was processed to evaluate the orbit determination consistency achieved using the Goddard Trajectory Determination System batch least-squares estimator. The effects of the use of the second operational relay spacecraft, of refinements in orbit determination models (geopotentials, polar motion, solid earth tidal gravitational perturbations, ionospheric refraction corrections), and of methods for providing relay spacecraft spacecraft position information were also studied.
On-orbit performance of Gravity Probe B orbit determination and drag-free translation control
NASA Astrophysics Data System (ADS)
Li, J.; Bencze, W. J.; Debra, D. B.; Galal, K.; Hanuschak, G.; Keiser, G. M.; Mester, J.; Shestople, P.; Small, H.
The Gravity Probe B GP-B Relativity Mission is a fundamental physics experiment to test Einstein s theory of General Relativity based on observations of gyros in a satellite in a 642 km circular polar orbit around the Earth The GP-B satellite is designed to test two predictions of Einstein s theory the geodetic effect and the frame-dragging effect to an extremely high accuracy Drag-free translation control is implemented to minimize support forces and support induced torques on the gyros One of the four redundant gyros is used as the proof mass and the propellant of the drag-free control system is derived from the exhaust gas boil-off from the helium dewar of the GP-B satellite The GP-B orbit is determined primarily from the measurements of the GPS receiver onboard the satellite and verified independently with the ground-based laser ranging data The force biases in both the attitude and translation control system and the gyro suspension system are also estimated in the ground processing of the orbit data and compensated in the drag-free control system This paper describes the design and implementation of the orbit determination and drag-free translation control system of the GP-B mission and shows the on-orbit performance from the launch on April 20 2004 to the depletion of the helium on September 29 2005
Orbit Determination for the 2007 Mars Phoenix Lander
NASA Technical Reports Server (NTRS)
Ryne, Mark S.; Graat, Eric; Haw, Robert; Kruizinga, Gerhard; Lau, Eunice; Martin-Mur, Tomas; McElrath, Timothy; Nandi, Sumita; Portock, Brian
2008-01-01
The Phoenix mission is designed to study the arctic region of Mars. To achieve this goal, the spacecraft must be delivered to a narrow corridor at the top of the Martian atmosphere, which is approximately 20 km wide. This paper will discuss the details of the Phoenix orbit determination process and the effort to reduce errors below the level necessary to achieve successful atmospheric entry at Mars. Emphasis will be placed on properly modeling forces that perturb the spacecraft trajectory and the errors and uncertainties associated with those forces. Orbit determination covariance analysis strongly influenced mission operations scenarios, which were chosen to minimize errors and associated uncertainties.
Evaluation of advanced geopotential models for operational orbit determination
NASA Technical Reports Server (NTRS)
Radomski, M. S.; Davis, B. E.; Samii, M. V.; Engel, C. J.; Doll, C. E.
1988-01-01
To meet future orbit determination accuracy requirements for different NASA projects, analyses are performed using Tracking and Data Relay Satellite System (TDRSS) tracking measurements and orbit determination improvements in areas such as the modeling of the Earth's gravitational field. Current operational requirements are satisfied using the Goddard Earth Model-9 (GEM-9) geopotential model with the harmonic expansion truncated at order and degree 21 (21-by-21). This study evaluates the performance of 36-by-36 geopotential models, such as the GEM-10B and Preliminary Goddard Solution-3117 (PGS-3117) models. The Earth Radiation Budget Satellite (ERBS) and LANDSAT-5 are the spacecraft considered in this study.
Filter Strategies for Mars Science Laboratory Orbit Determination
NASA Technical Reports Server (NTRS)
Thompson, Paul F.; Gustafson, Eric D.; Kruizinga, Gerhard L.; Martin-Mur, Tomas J.
2013-01-01
The Mars Science Laboratory (MSL) spacecraft had ambitious navigation delivery and knowledge accuracy requirements for landing inside Gale Crater. Confidence in the orbit determination (OD) solutions was increased by investigating numerous filter strategies for solving the orbit determination problem. We will discuss the strategy for the different types of variations: for example, data types, data weights, solar pressure model covariance, and estimating versus considering model parameters. This process generated a set of plausible OD solutions that were compared to the baseline OD strategy. Even implausible or unrealistic results were helpful in isolating sensitivities in the OD solutions to certain model parameterizations or data types.
5 CFR 1631.7 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 5 Administrative Personnel 3 2010-01-01 2010-01-01 false Initial determination. 1631.7 Section 1631.7 Administrative Personnel FEDERAL RETIREMENT THRIFT INVESTMENT BOARD AVAILABILITY OF RECORDS... determination. The FOIA Officer shall have the authority to approve or deny requests received pursuant to...
Optimal Control for a Cooperative Rendezvous Between Two Spacecraft from Determined Orbits
NASA Astrophysics Data System (ADS)
Feng, Weiming; Han, Liping; Shi, Lei; Zhao, Di; Yang, Kun
2016-01-01
The mathematical model of a far-distance cooperative rendezvous between two spacecraft in a non-Keplerian orbit was established. Approximate global optimization was performed by a type of hybrid algorithm consisting of particle swarm optimization and differential evolution. In this process, the double-fitness function was established according to the objective function and the constraints; the double-fitness function was used to enable a better choice between the solutions obtained by the two algorithms at every iteration. In addition, the costate variables obtained were set as the initial values of the sequential quadratic programming to greatly increase the possibility of finding the approximate global optimal solution. After performing the calculations and simulations, it was concluded that the fuel required for orbiting was not influenced by the initial positions of the two spacecraft if the initial orbits of the two spacecraft were determined. However, the time consumption is strongly influenced in this situation.
The role of laser determined orbits in geodesy and geophysics
NASA Technical Reports Server (NTRS)
Kolenkiewicz, R.; Smith, D. E.; Dunn, P. J.; Torrence, M. H.; Robbins, J. W.
1991-01-01
Some of the results of orbit analysis from the NASA SLR analysis group are presented. The earth's orientation was determined for 5-day intervals to 1.9 mas for the pole and 0.09 msec for length of day. The 3d center of mass station positions was determined to 33 mm over a period of 3 months, and geodesic rates of SLR tracking sites were determined to 5 mm/yr.
GRAIL Science Data System Orbit Determination : Approach, Strategy, and Performance
NASA Technical Reports Server (NTRS)
Fahnestock, Eugene; Asmar, Sami; Park, Ryan; Strekalov, Dmitry; Yuan, Dah-Ning; Harvey, Nate; Kahan, Daniel; Konopliv, Alex; Kruizinga, Gerhard; Oudrhiri, Kamal; Paik, Meegyeong
2013-01-01
This paper details orbit determination techniques and strategies employed within each stage of the larger iterative process of preprocessing raw GRAIL data into the gravity science measurements used within gravity field solutions. Each orbit determination pass used different data, corrections to them, and/or estimation parameters. We compare performance metrics among these passes. For example, for the primary mission, the magnitude of residuals using our orbits progressed from approximately or equal to19.4 to 0.077 approximately or equal to m/s for inter-satellite range rate data and from approximately or equal to 0.4 to approximately or equal to 0.1 mm/s for Doppler data.
(42355) Typhon Echidna: Scheduling observations for binary orbit determination
NASA Astrophysics Data System (ADS)
Grundy, W. M.; Noll, K. S.; Virtanen, J.; Muinonen, K.; Kern, S. D.; Stephens, D. C.; Stansberry, J. A.; Levison, H. F.; Spencer, J. R.
2008-09-01
We describe a strategy for scheduling astrometric observations to minimize the number required to determine the mutual orbits of binary transneptunian systems. The method is illustrated by application to Hubble Space Telescope observations of (42355) Typhon-Echidna, revealing that Typhon and Echidna orbit one another with a period of 18.971±0.006 days and a semimajor axis of 1628±29 km, implying a system mass of (9.49±0.52)×10 kg. The eccentricity of the orbit is 0.526±0.015. Combined with a radiometric size determined from Spitzer Space Telescope data and the assumption that Typhon and Echidna both have the same albedo, we estimate that their radii are 76-16+14 and 42-9+8 km, respectively. These numbers give an average bulk density of only 0.44-0.17+0.44 gcm, consistent with very low bulk densities recently reported for two other small transneptunian binaries.
Analysis of HY2A precise orbit determination using DORIS
NASA Astrophysics Data System (ADS)
Gao, Fan; Peng, Bibo; Zhang, Yu; Evariste, Ngatchou Heutchi; Liu, Jihua; Wang, Xiaohui; Zhong, Min; Lin, Mingsen; Wang, Nazi; Chen, Runjing; Xu, Houze
2015-03-01
HY2A is the first Chinese marine dynamic environment satellite. The payloads include a radar altimeter to measure the sea surface height in combination with a high precision orbit to be determined from tracking data. Onboard satellite tracking includes GPS, SLR, and the DORIS DGXX receiver which delivers phase and pseudo-range measurements. CNES releases raw phase and pseudo-range measurements with RINEX DORIS 3.0 format and pre-processed Doppler range-rate with DORIS 2.2 data format. However, the VMSI software package developed by Van Martin Systems, Inc which is used to estimate HY2A DORIS orbits can only process Doppler range-rate but not the DORIS phase data which are available with much shorter latency. We have proposed a method of constructing the phase increment data, which are similar to range-rate data, from RINEX DORIS 3.0 phase data. We compute the HY2A orbits from June, 2013 to August, 2013 using the POD strategy described in this paper based on DORIS 2.2 range-rate data and our reconstructed phase increment data. The estimated orbits are evaluated by comparing with the CNES precise orbits and SLR residuals. Our DORIS-only orbits agree with the precise GPS + SLR + DORIS CNES orbits radially at 1-cm and about 3-cm in the other two directions. SLR test with the 50° cutoff elevation shows that the CNES orbit can achieve about 1.1-cm accuracy in radial direction and our DORIS-only POD solutions are slightly worse. In addition, other HY2A DORIS POD concerns are discussed in this paper. Firstly, we discuss the frequency offset values provided with the RINEX data and find that orbit accuracy for the case when the frequency offset is applied is worse than when it is not applied. Secondly, HY2A DORIS antenna z-offsets are estimated using two kinds of measurements from June, 2013 to August, 2013. The results show that the measurement errors contribute a total of about 2-cm difference of estimated z-offset. Finally, we estimate HY2A orbits selecting 3 days with severe geomagnetic storm activity and SLR residuals suggest that estimating a drag coefficient every 6 h without any constraint is sufficient for maintaining orbit accuracy.
Orbit Determination for Mars Global Surveyor During Mapping
NASA Technical Reports Server (NTRS)
Lemoine, F. G.; Rowlands, D. D.; Smith, D. E.; Pavlis, D. E.; Chinn, D. S.; Luthcke, S. B.; Neumann, G. A.
1999-01-01
The Mars Global Surveyor (MGS) spacecraft reached a low-altitude circular orbit on February 4, 1999, after the termination of the second phase of aerobraking. The MGS spacecraft carries the Mars Orbiter Laser Altimeter (MOLA) whose primary goal is to derive a global, geodetically referenced 0.2 deg x 0.2 deg topographic grid of Mars with a vertical accuracy of better than 30 meters. During the interim science orbits in the' Hiatus mission phase (October - November 1997), and the Science Phasing Orbits (March - April, 1998, and June - July 1998) 208 passes of altimeter data were collected by the MOLA instrument. On March 1, 1999 the first ten orbits of MOLA altimeter data from the near-circular orbit were successfully returned from MGS by the Deep Space Network (DSN). Data will be collected from MOLA throughout the Mapping phase of the MCS mission, or for at least one Mars year (687 days). Whereas the interim orbits of Hiatus and SPO were highly eccentric, and altimeter data were only collected near periapsis when the spacecraft was below 785 km, the Mapping orbit of MGS is near circular, and altimeter data will be collected continuously at a rate of 10 Hz. The proper analysis of the altimeter data requires that the orbit of the MGS spacecraft be known to an accuracy comparable to that of the quality of the altimeter data. The altimeter has an ultimate precision of 30 cm on mostly flat surfaces, so ideally the orbits of the MGS spacecraft should be known to this level. This is a stringent requirement, and more realistic goals of orbit error for MGS are ten to thirty meters. In this paper we will discuss the force and measurement modelling required to achieve this objective. Issues in force modelling include the proper modelling of the gravity field of Mars, and the modelling of non-conservatives forces, including the development of a 'macro-model', in a similar fashion to TOPEX/POSEIDON and TDRSS. During Cruise and Aerobraking, the high gain antenna (HGA) was stowed on the +X face of the spacecraft. On March 29, 1999 the HGA will be deployed on a meter long boom which will remain Earth-pointing while the instrument panel (including the MOLA instrument) remains pointed at nadir. The tracking data must be corrected for the regular motion of the high gain antenna with respect to the center of mass, and the success of the MGS determination during Mapping will depend on correctly accounting for this offset in the measurement model.
Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter
NASA Technical Reports Server (NTRS)
Slojkowski, Steven; Lowe, Jonathan; Woodburn, James
2015-01-01
Orbit determination (OD) analysis results are presented for the Lunar Reconnaissance Orbiter (LRO) using a commercially available Extended Kalman Filter, Analytical Graphics' Orbit Determination Tool Kit (ODTK). Process noise models for lunar gravity and solar radiation pressure (SRP) are described and OD results employing the models are presented. Definitive accuracy using ODTK meets mission requirements and is better than that achieved using the operational LRO OD tool, the Goddard Trajectory Determination System (GTDS). Results demonstrate that a Vasicek stochastic model produces better estimates of the coefficient of solar radiation pressure than a Gauss-Markov model, and prediction accuracy using a Vasicek model meets mission requirements over the analysis span. Modeling the effect of antenna motion on range-rate tracking considerably improves residuals and filter-smoother consistency. Inclusion of off-axis SRP process noise and generalized process noise improves filter performance for both definitive and predicted accuracy. Definitive accuracy from the smoother is better than achieved using GTDS and is close to that achieved by precision OD methods used to generate definitive science orbits. Use of a multi-plate dynamic spacecraft area model with ODTK's force model plugin capability provides additional improvements in predicted accuracy.
An autonomous orbit determination method for MEO and LEO satellite
NASA Astrophysics Data System (ADS)
Zhang, Hui; Wang, Jin; Yu, Guobin; Zhong, Jie; Lin, Ling
2014-09-01
A reliable and secure navigation system and assured autonomous capability of satellite are in high demand in case of emergencies in space. This paper introduces a novel autonomous orbit determination method for Middle-Earth-Orbit and Low-Earth-Orbit (MEO and LEO) satellite by observing space objects whose orbits are known. Generally, the geodetic satellites, such as LAGEOS and ETALONS, can be selected as the space objects here. The precision CCD camera on tracking gimbal can make a series of photos of the objects and surrounding stars when MEO and LEO satellite encounters the space objects. Then the information processor processes images and attains sightings and angular observations of space objects. Several clusters of such angular observations are incorporated into a batch least squares filter to obtain an orbit determination solution. This paper describes basic principle and builds integrated mathematical model. The accuracy of this method is analyzed by means of computer simulation. Then a simulant experiment system is built, and the experimental results demonstrate the feasibility and effectiveness of this method. The experimental results show that this method can attain the accuracy of 150 meters with angular observations of 1 arcsecond system error.
Mars Science Laboratory Orbit Determination Data Pre-Processing
NASA Technical Reports Server (NTRS)
Gustafson, Eric D.; Kruizinga, Gerhard L.; Martin-Mur, Tomas J.
2013-01-01
The Mars Science Laboratory (MSL) was spin-stabilized during its cruise to Mars. We discuss the effects of spin on the radiometric data and how the orbit determination team dealt with them. Additionally, we will discuss the unplanned benefits of detailed spin modeling including attitude estimation and spacecraft clock correlation.
NASA Astrophysics Data System (ADS)
Tang, J. S.
2011-03-01
It has been over half a century since the launch of the first artificial satellite Sputnik in 1957, which marks the beginning of the Space Age. During the past 50 years, with the development and innovations in various fields and technologies, satellite application has grown more and more intensive and extensive. This thesis is based on three major research projects which the author joined in. These representative projects cover main aspects of satellite orbit theory and application of precise orbit determination (POD), and also show major research methods and important applications in orbit dynamics. Chapter 1 is an in-depth research on analytical theory of satellite orbits. This research utilizes general transformation theory to acquire high-order analytical solutions when mean-element method is not applicable. These solutions can be used in guidance and control or rapid orbit forecast within the accuracy of 10-6. We also discuss other major perturbations, each of which is considered with improved models, in pursuit of both convenience and accuracy especially when old models are hardly applicable. Chapter 2 is POD research based on observations. Assuming a priori force model and estimation algorithm have reached their accuracy limits, we introduce empirical forces to Shenzhou-type orbit in order to compensate possible unmodeled or mismodeled perturbations. Residuals are analyzed first and only empirical force models with actual physical background are considered. This not only enhances a posteriori POD accuracy, but also considerably improves the accuracy of orbit forecast. This chapter also contains theoretical discussions on modeling of empirical forces, computation of partial derivatives and propagation of various errors. Error propagation helps to better evaluate orbital accuracy in future missions. Chapter 3 is an application of POD in space geodesy. GRACE satellites are used to obtain Antarctic temporal gravity field between 2004 and 2007. Various changes from traditional methods are implemented to better represent the regional temporal gravity field in this work. As a thesis in astrodynamics, this chapter will concentrate on orbit problems and estimation approaches. Although most details in geophysics are skipped, gravity field solutions will be displayed and the preliminary images of Antarctic mass flux will be revealed. These researches are summarized but not concluded in this thesis. Many problems have been left in all the aspects mentioned in this thesis and need to be studied in future researches, not to mention that the fast developing space technology keeps redefining our traditional knowledge with new concepts and elements. So future work and directions will be discussed at the end of the thesis, expecting further progress upon the present achievements.
NASA Astrophysics Data System (ADS)
Yang, Yang; Yue, Xiaokui; Yuan, Jianping; Rizos, Chris
2014-11-01
Clock error estimation has been the focus of a great deal of research because of the extensive usage of clocks in GPS positioning applications. The receiver clock error in the spacecraft orbit determination is commonly estimated on an epoch-by-epoch basis, along with the spacecraft’s position. However, due to the high correlation between the spacecraft orbit altitude and the receiver clock parameters, estimates of the radial component are degraded in the kinematic approach. Using clocks with high stability, the predictable behaviour of the receiver oscillator can be exploited to improve the positioning accuracy, especially for the radial component. This paper introduces two GPS receiver clock models to describe the deterministic and stochastic property of the receiver clock, both of which can improve the accuracy of kinematic orbit determination for spacecraft in low earth orbit. In particular, the clock parameters are estimated as time offset and frequency offset in the two-state model. The frequency drift is also estimated as an unknown parameter in the three-state model. Additionally, residual non-deterministic random errors such as frequency white noise, frequency random walk noise and frequency random run noise are modelled. Test results indicate that the positioning accuracy could be improved significantly using one day of GRACE flight data. In particular, the error of the radial component was reduced by over 40.0% in the real-time scenario.
Expected orbit determination performance for the TOPEX/Poseidon mission
Nerem, R.S.; Putney, B.H.; Marshall, J.A.; Lerch, F.J. ); Pavlis, E.C. ); Klosko, S.M.; Luthcke, S.B.; Patel, G.B.; Williamson, R.G.; Zelensky, N.P.
1993-03-01
The TOPEX/Poseidon (T/P) mission, launched during the summer of 1992, has the requirement that the radial component of its orbit must be computed to an accuracy of 13 cm root-mean-square (rms) or better, allowing measurements of the sea surface height to be computed to similar accuracy when the satellite height is differenced with the altimeter measurements. This will be done by combining precise satellite tracking measurements with precise models of the forces acting on the satellite. The Space Geodesy Branch at Goddard Space Flight Center (GSFC), as part of the T/P precision orbit determination (POD) Team, has the responsibility within NASA for the T/P precise orbit computations. The prelaunch activities of the T/P POD Team have been mainly directed towards developing improved models of the static and time-varying gravitational forces acting on T/P and precise models for the non-conservative forces perturbing the orbit of T/P such as atmospheric drag, solar and Earth radiation pressure, and thermal imbalances. The radial orbit error budget for T/P allows 10 cm rms error due to gravity field mismodeling, 3 cm due to solid Earth and ocean tides, 6 cm due to radiative forces, and 3 cm due to atmospheric drag. A prelaunch assessment of the current modeling accuracies for these forces indicates that the radial orbit error requirements can be achieved with the current models, and can probably be surpassed once T/P tracking data are used to fine tune the models. Provided that the performance of the T/P spacecraft is nominal, the precise orbits computed by the T/P POD Team should be accurate to 13 cm or better radially.
GIOVE Orbit and Clock Determination Based on the CONGO Network
NASA Astrophysics Data System (ADS)
Steigenberger, Peter; Hauschild, AndrÃ©; Montenbruck, Oliver; Hugentobler, Urs; Hessels, Uwe; Weber, Georg; Noack, Thoralf
2010-05-01
As a prototype for the satellites of the future European Global Navigation Satellite System (GNSS) Galileo, the European Space Agency (ESA) launched two satellites (GIOVE-A and GIOVE-B) as part of the Galileo in Orbit Validation Element (GIOVE). To gain experience with the signals transmitted by these satellites and to estimate satellite orbit and clock parameters, a global network of GIOVE-capable receivers was established. This Cooperative Network for GIOVE Observations (CONGO) is operated by Deutsches Zentrum fÃ¼r Luft- und Raumfahrt (DLR, Oberpfaffenhofen, Germany) and Bundesamt fÃ¼r Kartographie und GeodÃ¤sie (BKG, Frankfurt, Germany) in cooperation with several local station hosts. The CONGO network currently consists of 10 globally distributed stations providing their observations in real-time. This network is used by Technische UniversitÃ¤t MÃ¼nchen for an operational daily orbit and clock determination of the GIOVE satellites including orbit predictions. The strategy of the combined GPS and GIOVE processing is presented. The quality of the estimated GIOVE satellite orbits is evaluated by orbit fits and satellite laser ranging (SLR). The quality of the GIOVE satellite clocks, in particular the hydrogen maser of GIOVE-B, is discussed. As three different receiver types and two different satellite systems are considered in the CONGO processing, a special focus has to be put on the biases between the different receivers and GNSSs. Additionally, DLR's Real-Time Clock Estimation (RETICLE) system has been extended to provide clock offset estimates for the GIOVE satellites based on the real-time data streams from the CONGO network. The GIOVE clocks are estimated based on the predicted orbits mentioned above. The paper introduces the real-time clock estimation process and presents real-time clock results.
Automated Orbit Determination System (AODS) requirements definition and analysis
NASA Technical Reports Server (NTRS)
Waligora, S. R.; Goorevich, C. E.; Teles, J.; Pajerski, R. S.
1980-01-01
The requirements definition for the prototype version of the automated orbit determination system (AODS) is presented including the AODS requirements at all levels, the functional model as determined through the structured analysis performed during requirements definition, and the results of the requirements analysis. Also specified are the implementation strategy for AODS and the AODS-required external support software system (ADEPT), input and output message formats, and procedures for modifying the requirements.
Orbit Determination of Hayabusa during Close Proximity Phase
NASA Astrophysics Data System (ADS)
Ikeda, Hitoshi; Kominato, Takashi; Matsuoka, Masatoshi; Ohnishi, Takafumi; Yoshikawa, Makoto
In September 2005, Hayabusa (MUSES-C) spacecraft successfully had a rendezvous with asteroid 25143 Itokawa. After the arrival, Hayabusa made detailed observations of the asteroid during its rendezvous period (about three months). As the results of various kinds of scientific analysis, a variety of physical parameters of Itokawa (e.g. size, volume, mass, and density) were derived. As to the orbit determination of Hayabusa spacecraft, during the cruise phase, the radiometric (2-way X-band range and Doppler) data were used for analysis. On the other hand, during the approach phase or rendezvous phase, we could obtain the optical data by means of star tracker or optical navigation camera, thus both the radiometric and the optical data were used for orbit determination. The present paper will report on the results of the orbit determination of Hayabusa during the close proximity phase. We will also mention about the mass estimation of Itokawa in this period. The data used in this analysis are 2-way X-band Doppler data and the position data, which were calculated from optical navigation camera's data. As well as the large orbital maneuvers and the gravitational acceleration of Itokawa, the effect of solar radiation pressure, and the effect of attitude control are also taken into account for the calculation. As to the gravity model of Itokawa, a spherical-harmonics gravity model or a polyhedron gravity model are used depending on the situation.
Hardware in-the-Loop Demonstration of Real-Time Orbit Determination in High Earth Orbits
NASA Technical Reports Server (NTRS)
Moreau, Michael; Naasz, Bo; Leitner, Jesse; Carpenter, J. Russell; Gaylor, Dave
2005-01-01
This paper presents results from a study conducted at Goddard Space Flight Center (GSFC) to assess the real-time orbit determination accuracy of GPS-based navigation in a number of different high Earth orbital regimes. Measurements collected from a GPS receiver (connected to a GPS radio frequency (RF) signal simulator) were processed in a navigation filter in real-time, and resulting errors in the estimated states were assessed. For the most challenging orbit simulated, a 12 hour Molniya orbit with an apogee of approximately 39,000 km, mean total position and velocity errors were approximately 7 meters and 3 mm/s respectively. The study also makes direct comparisons between the results from the above hardware in-the-loop tests and results obtained by processing GPS measurements generated from software simulations. Care was taken to use the same models and assumptions in the generation of both the real-time and software simulated measurements, in order that the real-time data could be used to help validate the assumptions and models used in the software simulations. The study makes use of the unique capabilities of the Formation Flying Test Bed at GSFC, which provides a capability to interface with different GPS receivers and to produce real-time, filtered orbit solutions even when less than four satellites are visible. The result is a powerful tool for assessing onboard navigation performance in a wide range of orbital regimes, and a test-bed for developing software and procedures for use in real spacecraft applications.
Orbit Determination Support for the Microwave Anisotropy Probe (MAP)
NASA Technical Reports Server (NTRS)
Bauer, Frank (Technical Monitor); Truong, Son H.; Cuevas, Osvaldo O.; Slojkowski, Steven
2003-01-01
NASA's Microwave Anisotropy Probe (MAP) was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle on June 30, 2001. The spacecraft received a nominal direct insertion by the Delta expendable launch vehicle into a 185-km circular orbit with a 28.7deg inclination. MAP was then maneuvered into a sequence of phasing loops designed to set up a lunar swingby (gravity-assisted acceleration) of the spacecraft onto a transfer trajectory to a lissajous orbit about the Earth-Sun L2 Lagrange point, about 1.5 million km from Earth. Because of its complex orbital characteristics, the mission provided a unique challenge for orbit determination (OD) support in many orbital regimes. This paper summarizes the premission trajectory covariance error analysis, as well as actual OD results. The use and impact of the various tracking stations, systems, and measurements will be also discussed. Important lessons learned from the MAP OD support team will be presented. There will be a discussion of the challenges presented to OD support including the effects of delta-Vs at apogee as well as perigee, and the impact of the spacecraft attitude mode on the OD accuracy and covariance analysis.
50 CFR 296.9 - Initial determination.
Code of Federal Regulations, 2013 CFR
2013-10-01
... 50 Wildlife and Fisheries 11 2013-10-01 2013-10-01 false Initial determination. 296.9 Section 296.9 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE CONTINENTAL SHELF FISHERMEN'S CONTINGENCY FUND § 296.9...
28 CFR 301.305 - Initial determination.
Code of Federal Regulations, 2012 CFR
2012-07-01
... delegated by the Board of Directors of Federal Prison Industries, Inc., pursuant to 28 CFR 0.99. In... 301.305 Judicial Administration FEDERAL PRISON INDUSTRIES, INC., DEPARTMENT OF JUSTICE INMATE ACCIDENT COMPENSATION Compensation for Work-Related Physical Impairment or Death § 301.305 Initial determination....
28 CFR 301.305 - Initial determination.
Code of Federal Regulations, 2014 CFR
2014-07-01
... delegated by the Board of Directors of Federal Prison Industries, Inc., pursuant to 28 CFR 0.99. In... 301.305 Judicial Administration FEDERAL PRISON INDUSTRIES, INC., DEPARTMENT OF JUSTICE INMATE ACCIDENT COMPENSATION Compensation for Work-Related Physical Impairment or Death § 301.305 Initial determination....
28 CFR 301.305 - Initial determination.
Code of Federal Regulations, 2011 CFR
2011-07-01
... delegated by the Board of Directors of Federal Prison Industries, Inc., pursuant to 28 CFR 0.99. In... 301.305 Judicial Administration FEDERAL PRISON INDUSTRIES, INC., DEPARTMENT OF JUSTICE INMATE ACCIDENT COMPENSATION Compensation for Work-Related Physical Impairment or Death § 301.305 Initial determination....
28 CFR 301.305 - Initial determination.
Code of Federal Regulations, 2013 CFR
2013-07-01
... delegated by the Board of Directors of Federal Prison Industries, Inc., pursuant to 28 CFR 0.99. In... 301.305 Judicial Administration FEDERAL PRISON INDUSTRIES, INC., DEPARTMENT OF JUSTICE INMATE ACCIDENT COMPENSATION Compensation for Work-Related Physical Impairment or Death § 301.305 Initial determination....
Orbit determination and orbit control for the Earth Observing System (EOS) AM spacecraft
NASA Technical Reports Server (NTRS)
Herberg, Joseph R.; Folta, David C.
1993-01-01
Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.
How to Determine an Exomoon's Sense of Orbital Motion
NASA Astrophysics Data System (ADS)
Heller, René; Albrecht, Simon
2014-11-01
We present two methods to determine an exomoon's sense of orbital motion (SOM), one with respect to the planet's circumstellar orbit and one with respect to the planetary rotation. Our simulations show that the required measurements will be possible with the European Extremely Large Telescope (E-ELT). The first method relies on mutual planet-moon events during stellar transits. Eclipses with the moon passing behind (in front of) the planet will be late (early) with regard to the moon's mean orbital period due to the finite speed of light. This "transit timing dichotomy" (TTD) determines an exomoon's SOM with respect to the circumstellar motion. For the 10 largest moons in the solar system, TTDs range between 2 and 12 s. The E-ELT will enable such measurements for Earth-sized moons around nearby Sun-like stars. The second method measures distortions in the IR spectrum of the rotating giant planet when it is transited by its moon. This Rossiter-McLaughlin effect (RME) in the planetary spectrum reveals the angle between the planetary equator and the moon's circumplanetary orbital plane, and therefore unveils the moon's SOM with respect to the planet's rotation. A reasonably large moon transiting a directly imaged planet like ? Pic b causes an RME amplitude of almost 100 m s-1, about twice the stellar RME amplitude of the transiting exoplanet HD209458 b. Both new methods can be used to probe the origin of exomoons, that is, whether they are regular or irregular in nature.
Magnetospheric Multiscale (MMS) Mission Commissioning Phase Orbit Determination Error Analysis
NASA Technical Reports Server (NTRS)
Chung, Lauren R.; Novak, Stefan; Long, Anne; Gramling, Cheryl
2009-01-01
The Magnetospheric MultiScale (MMS) mission commissioning phase starts in a 185 km altitude x 12 Earth radii (RE) injection orbit and lasts until the Phase 1 mission orbits and orientation to the Earth-Sun li ne are achieved. During a limited time period in the early part of co mmissioning, five maneuvers are performed to raise the perigee radius to 1.2 R E, with a maneuver every other apogee. The current baseline is for the Goddard Space Flight Center Flight Dynamics Facility to p rovide MMS orbit determination support during the early commissioning phase using all available two-way range and Doppler tracking from bo th the Deep Space Network and Space Network. This paper summarizes th e results from a linear covariance analysis to determine the type and amount of tracking data required to accurately estimate the spacecraf t state, plan each perigee raising maneuver, and support thruster cal ibration during this phase. The primary focus of this study is the na vigation accuracy required to plan the first and the final perigee ra ising maneuvers. Absolute and relative position and velocity error hi stories are generated for all cases and summarized in terms of the ma ximum root-sum-square consider and measurement noise error contributi ons over the definitive and predictive arcs and at discrete times inc luding the maneuver planning and execution times. Details of the meth odology, orbital characteristics, maneuver timeline, error models, and error sensitivities are provided.
Orbit determination based on meteor observations using numerical integration of equations of motion
NASA Astrophysics Data System (ADS)
Dmitriev, V.; Lupovka, V.; Gritsevich, M.
2014-07-01
We review the definitions and approaches to orbital-characteristics analysis applied to photographic or video ground-based observations of meteors. A number of camera networks dedicated to meteors registration were established all over the word, including USA, Canada, Central Europe, Australia, Spain, Finland and Poland. Many of these networks are currently operational. The meteor observations are conducted from different locations hosting the network stations. Each station is equipped with at least one camera for continuous monitoring of the firmament (except possible weather restrictions). For registered multi-station meteors, it is possible to accurately determine the direction and absolute value for the meteor velocity and thus obtain the topocentric radiant. Based on topocentric radiant one further determines the heliocentric meteor orbit. We aim to reduce total uncertainty in our orbit-determination technique, keeping it even less than the accuracy of observations. The additional corrections for the zenith attraction are widely in use and are implemented, for example, here [1]. We propose a technique for meteor-orbit determination with higher accuracy. We transform the topocentric radiant in inertial (J2000) coordinate system using the model recommended by IAU [2]. The main difference if compared to the existing orbit-determination techniques is integration of ordinary differential equations of motion instead of addition correction in visible velocity for zenith attraction. The attraction of the central body (the Sun), the perturbations by Earth, Moon and other planets of the Solar System, the Earth's flattening (important in the initial moment of integration, i.e. at the moment when a meteoroid enters the atmosphere), atmospheric drag may be optionally included in the equations. In addition, reverse integration of the same equations can be performed to analyze orbital evolution preceding to meteoroid's collision with Earth. To demonstrate the developed technique, we provide calculated orbits for several cases, including well-known meteorite-producing fireballs. A comparison of our estimates with previously published ones is also provided.
Orbital metastasis as initial manifestation of a widespread papillary thyroid microcarcinoma.
Pagsisihan, Daveric Ablis; Aguilar, Anthony Harvey Isabelo; Maningat, Ma Patricia Deanna Delfin
2015-01-01
Papillary thyroid carcinomas (PTCs), particularly microcarcinomas, rarely metastasise to the orbit. We report a case of a 49-year-old woman with a right supraorbital mass and unremarkable physical examination of the thyroid gland region. Orbital CT scan showed an expansile lytic lesion in the orbital plate of the frontal bone with a soft tissue component. An incision biopsy revealed metastatic well-differentiated thyroid carcinoma. Thyroid ultrasound was normal except for a subcentimetre nodule in the right lobe. The patient underwent total thyroidectomy where histopathology showed a subcentimetre follicular variant PTC. She subsequently received radioactive iodine therapy. Post-therapy whole body scan revealed metastatic thyroid tissues in the right orbital and posterior parietal, and left shoulder and hip areas. Although infrequent, metastatic thyroid carcinoma should be considered in patients with orbital metastasis even when neck examination is normal. In rare cases, this may be the initial manifestation of a widely metastatic papillary thyroid microcarcinoma. PMID:25819821
Orbit Determination Support for the Microwave Anisotropy Probe (MAP)
NASA Technical Reports Server (NTRS)
Truong, Son H.; Cuevas, Osvaldo O.; Slojkowski, Steven
2003-01-01
THe microwave Anisotropy Probe (MAP) ia the second Medium Class Explorer (MIDEX) mission of the National Aeronautics and Space Administration (NASA). The main goal of the MAP observatory is to measure the temperature fluctuations, known as anisotropy, of the cosmic microwave background (CBG) radiation over the entire sky and to produce a map of the CMB anisotropies with an angular resolution of approximately 3 degrees. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 20001. The spacecraft receiver a nominal direct insertion by the Delta expendable launch vehicle into a 185-km circular orbit with a 28.7 deg. inclination. MAP was than maneuvered into a sequence of phasing loops designed to set up a lunar gravity-assisted acceleration of the spacecraft onto a transfer trajectory to a lissajous orbit about the Earth-Sun L2 Lagrange point, about 1.5 million km from Earth. The science mission minimum lifetime is two years of observations at L2 with a desired lifetime of 4 years. The MAP transfer orbit consisted of 3.5 phasing loops.The MAP trajectory schematic all the way through L2 is shown. The first loop had a period of 7 days, the second and third loops were 10 days long, and the last half loop was 5 days. The periselene (i.e., lunar encounter or swingby) took place approximately 30 days after launch. After the periselene, the spacecraft cruised for approximately 60 days before it arrived in the vicinity of the L2 libration point. Two mid-course correction (MCC) maneuvers were performed to refine MAP's post-launch trajectory-one after periselene and one prior to arrival at vicinity of L2. Now that MAP is at its operational L2 lissajous orbit, the MAP satellite is commanded to perform occasional station-keeping (SK) maneuvers in order to maintain its orbit around L2. Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes.
Orbit Determination Support for the Microwave Anisotropy Probe (MAP)
NASA Technical Reports Server (NTRS)
Truong, Son H.; Cuevas, Osvaldo O.; Slojkowski, Steven; Bauer, Frank H. (Technical Monitor)
2002-01-01
The Microwave Anisotropy Probe (MAP) is a Medium Class Explorers (MIDEX) mission produced in partnership between Goddard Space Flight Center (GSFC) and Princeton University. The main science objective of the MAP mission is to produce an accurate full-sky map of the cosmic microwave background temperature fluctuations anisotropy. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 2001. The spacecraft received a nominal direct insertion by the Delta into a 185 km circular orbit. MAP was then maneuvered into a sequence of phasing loops designed to set up a lunar swingby (gravity-assisted acceleration) of the spacecraft onto a transfer trajectory to a Lissajous orbit about the Earth-Sun L2 point. The mission duration is approximately 27 months with 3 to 4 months of transfer time to the final mission orbit about L2. The MAP transfer orbit consisted of 3.5 phasing loops: the first loop has a 7-day period, the second and third loops have a 9-day period, and the last half loop has a 4-day period as illustrated in Figure 1, which also indicates the placement of maneuvers. A Pfinal correction maneuver was performed 18 hours after the last perigee to more closely achieve the targeted lissajous orbit. The lunar encounter or swingby took place approximately 30 days after launch. After the lunar encounter, the spacecraft will cruise for approximately 120 days before it arrives at L2. A Mid-Course Correction (MCC) maneuver was executed seven days after the swingby to further refine the trajectory. Once the MAP satellite is injected into the L2 Lissajous orbit, it will perform occasional stationkeeping maneuvers to maintain the Lissajous orbit for a minimum of two years (and a goal of four years). Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes. Extensive trajectory error covariance analysis was performed to predict ephemeris accuracy for the OD process using a Bayesian least-squares technique. The orbit determination error analysis is essential for maneuver planning and maneuver recovery study. Several tracking scenarios were investigated for each phase of the mission. This paper provides a summary of the premission trajectory covariance error analysis, as well as actual real-time OD results. The use and impact of the various tracking stations, systems, and measurements will be discussed. Details of the operational OD support and the inferred OD accuracy will be presented, and the results will be compared to the premission covariance analysis, In addition, there will be a discussion of the challenges presented to OD support including delta-Vs at apogee as well as perigee, and effects due to spacecraft attitude mode, in light of their implications to the OD accuracy and covariance analysis.
Operational Challenges In TDRS Post-Maneuver Orbit Determination
NASA Technical Reports Server (NTRS)
Laing, Jason; Myers, Jessica; Ward, Douglas; Lamb, Rivers
2015-01-01
The GSFC Flight Dynamics Facility (FDF) is responsible for daily and post maneuver orbit determination for the Tracking and Data Relay Satellite System (TDRSS). The most stringent requirement for this orbit determination is 75 meters total position accuracy (3-sigma) predicted over one day for Terra's onboard navigation system. To maintain an accurate solution onboard Terra, a solution is generated and provided by the FDF Four hours after a TDRS maneuver. A number of factors present challenges to this support, such as maneuver prediction uncertainty and potentially unreliable tracking from User satellities. Reliable support is provided by comparing an extended Kalman Filter (estimated using ODTK) against a Batch Least Squares system (estimated using GTDS).
Meteoroid and Orbital Debris Threats to NASA's Docking Seals: Initial Assessment and Methodology
NASA Technical Reports Server (NTRS)
deGroh, Henry C., III; Nahra, Henry K.
2009-01-01
The Crew Exploration Vehicle (CEV) will be exposed to the Micrometeoroid Orbital Debris (MMOD) environment in Low Earth Orbit (LEO) during missions to the International Space Station (ISS) and to the micrometeoroid environment during lunar missions. The CEV will be equipped with a docking system which enables it to connect to ISS and the lunar module known as Altair; this docking system includes a hatch that opens so crew and supplies can pass between the spacecrafts. This docking system is known as the Low Impact Docking System (LIDS) and uses a silicone rubber seal to seal in cabin air. The rubber seal on LIDS presses against a metal flange on ISS (or Altair). All of these mating surfaces are exposed to the space environment prior to docking. The effects of atomic oxygen, ultraviolet and ionizing radiation, and MMOD have been estimated using ground based facilities. This work presents an initial methodology to predict meteoroid and orbital debris threats to candidate docking seals being considered for LIDS. The methodology integrates the results of ground based hypervelocity impacts on silicone rubber seals and aluminum sheets, risk assessments of the MMOD environment for a variety of mission scenarios, and candidate failure criteria. The experimental effort that addressed the effects of projectile incidence angle, speed, mass, and density, relations between projectile size and resulting crater size, and relations between crater size and the leak rate of candidate seals has culminated in a definition of the seal/flange failure criteria. The risk assessment performed with the BUMPER code used the failure criteria to determine the probability of failure of the seal/flange system and compared the risk to the allotted risk dictated by NASA's program requirements.
Meteoroid and Orbital Debris Threats to NASA's Docking Seals: Initial Assessment and Methodology
NASA Technical Reports Server (NTRS)
deGroh, Henry C., III; Gallo, Christopher A.; Nahra, Henry K.
2009-01-01
The Crew Exploration Vehicle (CEV) will be exposed to the Micrometeoroid Orbital Debris (MMOD) environment in Low Earth Orbit (LEO) during missions to the International Space Station (ISS) and to the micrometeoroid environment during lunar missions. The CEV will be equipped with a docking system which enables it to connect to ISS and the lunar module known as Altair; this docking system includes a hatch that opens so crew and supplies can pass between the spacecrafts. This docking system is known as the Low Impact Docking System (LIDS) and uses a silicone rubber seal to seal in cabin air. The rubber seal on LIDS presses against a metal flange on ISS (or Altair). All of these mating surfaces are exposed to the space environment prior to docking. The effects of atomic oxygen, ultraviolet and ionizing radiation, and MMOD have been estimated using ground based facilities. This work presents an initial methodology to predict meteoroid and orbital debris threats to candidate docking seals being considered for LIDS. The methodology integrates the results of ground based hypervelocity impacts on silicone rubber seals and aluminum sheets, risk assessments of the MMOD environment for a variety of mission scenarios, and candidate failure criteria. The experimental effort that addressed the effects of projectile incidence angle, speed, mass, and density, relations between projectile size and resulting crater size, and relations between crater size and the leak rate of candidate seals has culminated in a definition of the seal/flange failure criteria. The risk assessment performed with the BUMPER code used the failure criteria to determine the probability of failure of the seal/flange system and compared the risk to the allotted risk dictated by NASA s program requirements.
Improved DORIS accuracy for precise orbit determination and geodesy
NASA Technical Reports Server (NTRS)
Willis, Pascal; Jayles, Christian; Tavernier, Gilles
2004-01-01
In 2001 and 2002, 3 more DORIS satellites were launched. Since then, all DORIS results have been significantly improved. For precise orbit determination, 20 cm are now available in real-time with DIODE and 1.5 to 2 cm in post-processing. For geodesy, 1 cm precision can now be achieved regularly every week, making now DORIS an active part of a Global Observing System for Geodesy through the IDS.
NASA Astrophysics Data System (ADS)
Bennett, J.; Sang, J.; Smith, C.; Zhang, K.
2014-09-01
In this paper results are presented from a short-arc orbit determination study using optical and laser tracking data from the Space Debris Tracking System located at Mount Stromlo, Australia. Fifteen low-Earth orbit debris objects were considered in the study with perigee altitudes in the range 550850 km. In most cases, a 2-day orbit determination was considered using 2 passes of optical and 2 passes of laser tracking data. A total of 33 1-day and 26 2-day orbit prediction cases were compared with residuals obtained by comparing the orbit prediction with subsequent tracking data. A comparison was made between the orbit prediction accuracies for 2 orbit determination variants: (1) Entire passes are used during the orbit determination process; (2) Only 5 seconds is used from the beginning of each pass. Overall, the short-arc orbit determination results in (slightly) worse 1 and 2 day orbit prediction accuracies when compared to using the full observation arcs; however, the savings in tracking load outweighs the reduction in accuracy. If the optical or laser data is left out of the 5-second pass orbit determination process, most cases diverged which shows the importance of 3-dimenional positioning. Two-line element data was used to constrain the orbit determination process resulting in better convergence rates, but the resulting orbit prediction accuracy was much worse. The results have important implications for an optical and laser debris tracking network with potential savings in tracking load. An experimental study will be needed to verify this statement.
Interpolation schemes for orbit determination with the global positioning system
NASA Technical Reports Server (NTRS)
Argentiero, P.; Morduch, G. E.
1977-01-01
This paper demonstrates that the Global Positioning System (GPS) and simple interpolation schemes can be utilized to satisfy typical orbit determination demands of applications satellites. The complete GPS consists of 24 satellites and permits a position fix of a user satellite at any arbitrary instant. The interpolation formulae used in this report fit generalized Keplerian orbits through a number of position fixes. For a given accuracy level the telemetry requirements as measured by the time intervals between position fixes and the computational load of an interpolation as measured by the number of points through which a generalized Keplerian orbit is fitted vary inversely. A set of possible compromises between these two factors is presented. The Phase I GPS consists of 6 satellites and permits a position fix of a user satellite just 36% of the time. It is shown that with this system more sophisticated interpolation schemes which model atmospheric drag and higher degree terms of the earth's gravity field must be employed to obtain accurate orbits.
Galileo satellites measurement biases and orbit determination: first results
NASA Astrophysics Data System (ADS)
Perosanz, F.; Loyer, S.; Mercier, F.; Boulanger, C.; Capdeville, H.; Mezerette, A.
2012-12-01
Thanks to the IGS Multi-GNSS Experiment (M-GEX), signals from new GNSS satellites like Galileo are now available. CNES and IGN joined their efforts to contribute to the densification of this multi-GNSS global network through the REGINA project. However this network includes geodetic receivers from several manufacturers. For this reason we realized a dedicated test campaign to characterize the different receivers available in order to be able to process in a consistent way the data from the MGEX network. The test consisted in zero baseline measurements between receivers. Pseudo range as well as phase and wide-lane biases have been identified between Trimble, Leica, Javad and Septentrio receivers. Then the data from the global M-GEX tracking network have been processed for the Precise Orbit determination of the Galileo satellite. The strategy followed the one that the CNES-CLS IGS Analysis Center uses to compute hybrid GPS-GLONASS products. Since July 2012, Galileo data are processed and orbit solutions are routinely produced and evaluated. Pseudo-range and phase biases between receiver as well as inter-system biases have been quantified. We also demonstrated that a decimeter 3D-WRMS orbit accuracy of Galileo satellite orbit can be achieved even during the constellation deployment.
Galileo satellites measurement biases and orbit determination : preliminary results
NASA Astrophysics Data System (ADS)
Perosanz, Felix; Loyer, Sylvain; Mercier, Flavien; Boulanger, Cyrille; Capdeville, Hugues; Mezerette, Adrien
2013-04-01
Thanks to the IGS Multi-GNSS Experiment (M-GEX), signals from new GNSS satellites like Galileo are now available. CNES and IGN joined their efforts to contribute to the densification of this multi-GNSS global network through the REGINA project. However this network includes geodetic receivers from several manufacturers. For this reason we realized a dedicated test campaign to characterize the different receivers available in order to be able to process in a consistent way the data from the MGEX network. The test consisted in zero baseline measurements between receivers. Pseudo range as well as phase and wide-lane biases have been identified between Trimble, Leica, Javad and Septentrio receivers. Then the data from the global M-GEX tracking network have been processed for the Precise Orbit determination (POD) of the Galileo satellite. The strategy followed the one that the CNES-CLS IGS Analysis Center uses to compute hybrid GPS-GLONASS products. Since July 2012, Galileo data are processed and orbit solutions are routinely produced and evaluated. Pseudo-range and phase biases between receiver as well as inter-system biases have been quantified. We also demonstrated that a sub-decimeter 3D-WRMS orbit accuracy of Galileo satellite orbit can be achieved even during the constellation deployment.
Gravity Recovery and Interior Laboratory Mission (GRAIL) Orbit Determination
NASA Technical Reports Server (NTRS)
You, Tung-Han; Antreasian, Peter; Broschart, Stephen; Criddle, Kevin; Higa, Earl; Jefferson, David; Lau, Eunice; Mohan, Swati; Ryne, Mark; Keck, Mason
2012-01-01
Launched on 10 September 2011 from the Cape Canaveral Air Force Station, Florida, the twin-spacecraft Gravity Recovery and Interior Laboratory (GRAIL) has the primary mission objective of generating a lunar gravity map with an unprecedented resolution via the Ka-band Lunar Gravity Ranging System (LGRS). After successfully executing nearly 30 maneuvers on their six-month journey, Ebb and Flow (aka GRAIL-A and GRAIL-B) established the most stringent planetary formation orbit on 1 March 2012 of approximately 30 km x 90 km in orbit size. This paper describes the orbit determination (OD) filter configurations, analyses, and results during the Trans-Lunar Cruise, Orbit Period Reduction, and Transition to Science Formation phases. The maneuver reconstruction strategies and their performance will also be discussed, as well as the navigation requirements, major dynamic models, and navigation challenges. GRAIL is the first mission to generate a full high-resolution gravity field of the only natural satellite of the Earth. It not only enables scientists to understand the detailed structure of the Moon but also further extends their knowledge of the evolutionary histories of the rocky inner planets. Robust and successful navigation was the key to making this a reality.
Enhanced orbit determination filter sensitivity analysis: Error budget development
NASA Technical Reports Server (NTRS)
Estefan, J. A.; Burkhart, P. D.
1994-01-01
An error budget analysis is presented which quantifies the effects of different error sources in the orbit determination process when the enhanced orbit determination filter, recently developed, is used to reduce radio metric data. The enhanced filter strategy differs from more traditional filtering methods in that nearly all of the principal ground system calibration errors affecting the data are represented as filter parameters. Error budget computations were performed for a Mars Observer interplanetary cruise scenario for cases in which only X-band (8.4-GHz) Doppler data were used to determine the spacecraft's orbit, X-band ranging data were used exclusively, and a combined set in which the ranging data were used in addition to the Doppler data. In all three cases, the filter model was assumed to be a correct representation of the physical world. Random nongravitational accelerations were found to be the largest source of error contributing to the individual error budgets. Other significant contributors, depending on the data strategy used, were solar-radiation pressure coefficient uncertainty, random earth-orientation calibration errors, and Deep Space Network (DSN) station location uncertainty.
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.; Zelensky, Nikita P.; Rowlands, David D.; Lemoine, Frank G.; Williams, Teresa A.
2003-01-01
Jason-1, launched on December 7, 2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. Jason-1 is no exception and has set a 1 cm radial orbit accuracy goal, which represents a factor of two improvement over what is currently being achieved for T/P. The challenge to precision orbit determination (POD) is both achieving the 1 cm radial orbit accuracy and evaluating and validating the performance of the 1 cm orbit. Fortunately, Jason-1 POD can rely on four independent tracking data types including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. In addition, to the enhanced GPS receiver, Jason-1 carries significantly improved SLR and DORIS tracking systems along with the altimeter itself. We demonstrate the 1 cm radial orbit accuracy goal has been achieved using GPS data alone in a reduced dynamic solution. It is also shown that adding SLR data to the GPS-based solutions improves the orbits even further. In order to assess the performance of these orbits it is necessary to process all of the available tracking data (GPS, SLR, DORIS and altimeter crossover differences) as either dependent or independent of the orbit solutions. It was also necessary to compute orbit solutions using various combinations of the four available tracking data in order to independently assess the orbit performance. Towards this end, we have greatly improved orbits determined solely from SLR+DORIS data by applying the reduced dynamic solution strategy. In addition, we have computed reduced dynamic orbits based on SLR, DORIS and crossover data that are a significant improvement over the SLR and DORIS based dynamic solutions. These solutions provide the best performing orbits for independent validation of the GPS-based reduced dynamic orbits.
GPS-Based Navigation And Orbit Determination for the AMSAT AO-40 Satellite
NASA Technical Reports Server (NTRS)
Davis, George; Moreau, Michael; Carpenter, Russell; Bauer, Frank
2002-01-01
The AMSAT OSCAR-40 (AO-40) spacecraft occupies a highly elliptical orbit (HEO) to support amateur radio experiments. An interesting aspect of the mission is the attempted use of GPS for navigation and attitude determination in HEO. Previous experiences with GPS tracking in such orbits have demonstrated the ability to acquire GPS signals, but very little data were produced for navigation and orbit determination studies. The AO-40 spacecraft, flying two Trimble Advanced Navigation Sensor (TANS) Vector GPS receivers for signal reception at apogee and at perigee, is the first to demonstrate autonomous tracking of GPS signals from within a HEO with no interaction from ground controllers. Moreover, over 11 weeks of total operations as of June 2002, the receiver has returned a continuous stream of code phase, Doppler, and carrier phase measurements useful for studying GPS signal characteristics and performing post-processed orbit determination studies in HEO. This paper presents the initial efforts to generate AO-40 navigation solutions from pseudorange data reconstructed from the TANS Vector code phase, as well as to generate a precise orbit solution for the AO-40 spacecraft using a batch filter.
CODE's new solar radiation pressure model for GNSS orbit determination
NASA Astrophysics Data System (ADS)
Arnold, D.; Meindl, M.; Beutler, G.; Dach, R.; Schaer, S.; Lutz, S.; Prange, L.; SoÅ›nica, K.; Mervart, L.; JÃ¤ggi, A.
2015-08-01
The Empirical CODE Orbit Model (ECOM) of the Center for Orbit Determination in Europe (CODE), which was developed in the early 1990s, is widely used in the International GNSS Service (IGS) community. For a rather long time, spurious spectral lines are known to exist in geophysical parameters, in particular in the Earth Rotation Parameters (ERPs) and in the estimated geocenter coordinates, which could recently be attributed to the ECOM. These effects grew creepingly with the increasing influence of the GLONASS system in recent years in the CODE analysis, which is based on a rigorous combination of GPS and GLONASS since May 2003. In a first step we show that the problems associated with the ECOM are to the largest extent caused by the GLONASS, which was reaching full deployment by the end of 2011. GPS-only, GLONASS-only, and combined GPS/GLONASS solutions using the observations in the years 2009-2011 of a global network of 92 combined GPS/GLONASS receivers were analyzed for this purpose. In a second step we review direct solar radiation pressure (SRP) models for GNSS satellites. We demonstrate that only even-order short-period harmonic perturbations acting along the direction Sun-satellite occur for GPS and GLONASS satellites, and only odd-order perturbations acting along the direction perpendicular to both, the vector Sun-satellite and the spacecraft's solar panel axis. Based on this insight we assess in the third step the performance of four candidate orbit models for the future ECOM. The geocenter coordinates, the ERP differences w. r. t. the IERS 08 C04 series of ERPs, the misclosures for the midnight epochs of the daily orbital arcs, and scale parameters of Helmert transformations for station coordinates serve as quality criteria. The old and updated ECOM are validated in addition with satellite laser ranging (SLR) observations and by comparing the orbits to those of the IGS and other analysis centers. Based on all tests, we present a new extended ECOM which substantially reduces the spurious signals in the geocenter coordinate (by about a factor of 2-6), reduces the orbit misclosures at the day boundaries by about 10 %, slightly improves the consistency of the estimated ERPs with those of the IERS 08 C04 Earth rotation series, and substantially reduces the systematics in the SLR validation of the GNSS orbits.
Orbit determination support for Hiten's aerobraking in the Earth's atmosphere
NASA Astrophysics Data System (ADS)
Efron, L.; Ellis, J.; Menon, P. R.; Tucker, B.
1991-12-01
Two passes of the ISAS (Japan's Institute of Space and Astronautical Science) Hiten spacecraft through the Earth's atmosphere, at perigee altitudes of 125 km and 120 km, during Mar. 1991 marked the first aerobraking technology demonstrations for an object in cis-lunar orbit traveling at near Earth escape velocity. Prediction and control of perigee altitude to better than 1 km was desired to assure spacecraft survival. Covariance analysis provided confidence that prediction accuracy better than 200 m for support of final trim maneuver design was achievable with NASA DSN (Deep Space Network) tracking support. ISAS used orbit determination results in their decisions to cancel final trim maneuvers. Post flight reconstructions, marking the first combined use of DSN and ISAS tracking data, yielded perigee altitude solutions which agree with the near real time mission operations support predictions to better than 50 m.
HOW TO DETERMINE AN EXOMOON'S SENSE OF ORBITAL MOTION
Heller, RenÃ©; Albrecht, Simon E-mail: albrecht@phys.au.dk
2014-11-20
We present two methods to determine an exomoon's sense of orbital motion (SOM), one with respect to the planet's circumstellar orbit and one with respect to the planetary rotation. Our simulations show that the required measurements will be possible with the European Extremely Large Telescope (E-ELT). The first method relies on mutual planet-moon events during stellar transits. Eclipses with the moon passing behind (in front of) the planet will be late (early) with regard to the moon's mean orbital period due to the finite speed of light. This ''transit timing dichotomy'' (TTD) determines an exomoon's SOM with respect to the circumstellar motion. For the 10 largest moons in the solar system, TTDs range between 2 and 12 s. The E-ELT will enable such measurements for Earth-sized moons around nearby Sun-like stars. The second method measures distortions in the IR spectrum of the rotating giant planet when it is transited by its moon. This Rossiter-McLaughlin effect (RME) in the planetary spectrum reveals the angle between the planetary equator and the moon's circumplanetary orbital plane, and therefore unveils the moon's SOM with respect to the planet's rotation. A reasonably large moon transiting a directly imaged planet like Î²Â PicÂ b causes an RME amplitude of almost 100 m s{sup â€“1}, about twice the stellar RME amplitude of the transiting exoplanet HD209458Â b. Both new methods can be used to probe the origin of exomoons, that is, whether they are regular or irregular in nature.
An independent determination of Fomalhaut b's orbit and the dynamical effects on the outer dust belt
NASA Astrophysics Data System (ADS)
Beust, H.; Augereau, J.-C.; Bonsor, A.; Graham, J. R.; Kalas, P.; Lebreton, J.; Lagrange, A.-M.; Ertel, S.; Faramaz, V.; ThÃ©bault, P.
2014-01-01
Context. The nearby star Fomalhaut harbors a cold, moderately eccentric (e ~ 0.1) dust belt with a sharp inner edge near 133 au. A low-mass, common proper motion companion, Fomalhaut b (Fom b), was discovered near the inner edge and was identified as a planet candidate that could account for the belt morphology. However, the most recent orbit determination based on four epochs of astrometry over eight years reveals a highly eccentric orbit (e = 0.8 Â± 0.1) that appears to cross the belt in the sky plane projection. Aims: We perform here a full orbital determination based on the available astrometric data to independently validate the orbit estimates previously presented. Adopting our values for the orbital elements and their associated uncertainties, we then study the dynamical interaction between the planet and the dust ring, to check whether the proposed disk sculpting scenario by Fom b is plausible. Methods: We used a dedicated MCMC code to derive the statistical distributions of the orbital elements of Fom b. Then we used symplectic N-body integration to investigate the dynamics of the dust belt, as perturbed by a single planet. Different attempts were made assuming different masses for Fom b. We also performed a semi-analytical study to explain our results. Results: Our results are in good agreement with others regarding the orbit of Fom b. We find that the orbit is highly eccentric, is close to apsidally aligned with the belt, and has a mutual inclination relative to the belt plane of <29Â° (67% confidence). If coplanar, this orbit crosses the disk. Our dynamical study then reveals that the observed planet could sculpt a transient belt configuration with a similar eccentricity to what is observed, but it would not be simultaneously apsidally aligned with the planet. This transient configuration only occurs a short time after the planet is placed on such an orbit (assuming an initially circular disk), a time that is inversely proportional to the planet's mass, and that is in any case much less than the 440 Myr age of the star. Conclusions: We constrain how long the observed dust belt could have survived with Fom b on its current orbit, as a function of its possible mass. This analysis leads us to conclude that Fom b is likely to have low mass, that it is unlikely to be responsible for the sculpting of the belt, and that it supports the hypothesis of a more massive, less eccentric planet companion Fomalhaut c.
Improving GLONASS Precise Orbit Determination through Data Connection
Liu, Yang; Ge, Maorong; Shi, Chuang; Lou, Yidong; Wickert, Jens; Schuh, Harald
2015-01-01
In order to improve the precision of GLONASS orbits, this paper presents a method to connect the data segments of a single station-satellite pair to increase the observation continuity and, consequently, the strength of the precise orbit determination (POD) solution. In this method, for each GLONASS station-satellite pair, the wide-lane ambiguities derived from the Melbourneâ€“WÃ¼bbena combination are statistically tested and corrected for phase integer offsets and then the same is carried out for the narrow-lane ambiguities calculated from the POD solution. An experimental validation was carried out using one-month GNSS data of a global network with 175 IGS stations. The result shows that, on average, 27.1% of the GLONASS station-satellite pairs with multiple data segments could be connected to a single long observation arc and, thus, only one ambiguity parameter was estimated. Using the connected data, the GLONASS orbit overlapping RMS at the day boundaries could be reduced by 19.2% in ideal cases with an averaged reduction of about 6.3%. PMID:26633414
Improving GLONASS Precise Orbit Determination through Data Connection.
Liu, Yang; Ge, Maorong; Shi, Chuang; Lou, Yidong; Wickert, Jens; Schuh, Harald
2015-01-01
In order to improve the precision of GLONASS orbits, this paper presents a method to connect the data segments of a single station-satellite pair to increase the observation continuity and, consequently, the strength of the precise orbit determination (POD) solution. In this method, for each GLONASS station-satellite pair, the wide-lane ambiguities derived from the Melbourne-WÃ¼bbena combination are statistically tested and corrected for phase integer offsets and then the same is carried out for the narrow-lane ambiguities calculated from the POD solution. An experimental validation was carried out using one-month GNSS data of a global network with 175 IGS stations. The result shows that, on average, 27.1% of the GLONASS station-satellite pairs with multiple data segments could be connected to a single long observation arc and, thus, only one ambiguity parameter was estimated. Using the connected data, the GLONASS orbit overlapping RMS at the day boundaries could be reduced by 19.2% in ideal cases with an averaged reduction of about 6.3%. PMID:26633414
Improved initialization conditions and single impulsive maneuvers for -invariant relative orbits
NASA Astrophysics Data System (ADS)
Dang, Zhaohui; Wang, Zhaokui; Zhang, Yulin
2015-03-01
The determination of the initial conditions for long-term bounded relative motion under natural perturbations is an important theme in satellite cluster flight. Considering the most significant perturbation of the geopotential, namely, the term, many researchers have proposed -mitigating initial conditions for satellite-bounded relative motion. To improve the existing -invariant conditions, two new methods for finding the correction factor are presented in this paper. In these two methods, Method 1 is obtained by minimizing the possible maximum drift in the along-track relative motion. However, Method 2 is designed by nullifying the rates of change of the bounds of the relative motion. Then the geometric character, such as the self-intersection of the -invariant relative orbits, is discussed. The conditions of 0, 1 and 2 (the possible maximum number) self-intersection points are also derived. Then, using Gauss's equations of planetary motion, an analytical optimal single-impulsive maneuver is deduced to guarantee the secular bounded relative motion under , too. Some numerical simulations are performed to validate the corresponding theoretical predictions. The results demonstrate that the proposed methods enhance performance for achieving the bounded relative motion under effects and can be used for future satellite cluster flight missions.
Galileo orbit determination from launch through the first earth flyby
NASA Technical Reports Server (NTRS)
Pollmeier, V. M.; Kallemeyn, P. H.
1991-01-01
The data types used in the Galileo orbit determination process and the primary effects on the spacecraft for the first two trajectory legs of the mission are discussed. These types are: two-way coherent Doppler, two-way range, and delta differential one-way range. Attention is given to four primary nongravitational accelerations that affect the Galileo spacecraft: Delta(V) from TCMs, the Delta(V) from attitude updates, the Delta(V) from thruster maintenance events, and solar radiation pressure.
Application of semianalytical satellite theories to precision orbit determination
NASA Technical Reports Server (NTRS)
Cefola, P. J.
1978-01-01
Those factors which limit the usefulness of current implementations of the semianalytical approach are discussed. Numerical and analytical enhancements to the semianalytical approach are considered. A simple mathematical model is provided to estimate the computational speed of a semianalytical theory employing the suggested enhancements. The model can factor in current experience with semianalytical theories (integration stepsizes, quadrature orders, speed of recursive formulations, etc.) and the characteristics of the particular output requirement (observation span (or orbit determination interval), observation rate, observation model, etc.). Comparisons with numerical integration are suggested.
Radiation force modeling for ICESat precision orbit determination
NASA Astrophysics Data System (ADS)
Webb, Charles Edward
2007-12-01
Precision orbit determination (POD) for the Ice, Cloud and land Elevation Satellite (ICESat) relies on an epoch-state batch filter, in which the dynamic models play a central role. Its implementation in the Multi-Satellite Orbit Determination Program (MSODP) originally included a box-and-wing model, representing the TOPEX/Poseidon satellite, to compute solar radiation forces. This "macro-model" has been adapted to the ICESat geometry, and additionally, extended to the calculation of forces induced by radiation reflected and emitted from the Earth. To determine the area and reflectivity parameters of the ICESat macro-model surfaces, a high-fidelity simulation of the radiation forces in low-Earth orbit was first developed, using a detailed model of the satellite, called the "micro-model". In this effort, new algorithms to compute such forces were adapted from a Monte Carlo Ray Tracing (MCRT) method originally designed to determine incident heating rates. After working with the vendor of the Thermal Synthesizer System (TSS) to implement these algorithms, a modified version of this software was employed to generate solar and Earth radiation forces for all ICESat orbit and attitude geometries. Estimates of the macro-model parameters were then obtained from a least-squares fit to these micro-model forces, applying an algorithm that also incorporated linear equality and inequality constraints to ensure feasible solutions. Three of these fitted solutions were selected for post-launch evaluation. Two represented conditions at the start and at the end of the mission, while the third comprised four separate solutions, one for each of the nominal satellite attitudes. In addition, three other sets of macro-model parameters were derived from area-weighted averaging of the micro-model reflectivities. They included solar-only and infrared-only spectral parameters, as well as a set combining these parameters. Daily POD solutions were generated with each of these macro-model sets, for eight-day intervals in four different ICESat mapping campaigns. As a group, the fitted parameters slightly outperformed the averaged parameters, based on a variety of metrics. Their impact on POD accuracy, however, was limited to the sub-millimeter level, as measured by independent satellite laser ranging (SLR) residuals. As a result, no change to the nominal macro-model parameters is recommended.
Orbital period determination in an eclipsing dwarf nova HT Cas
NASA Astrophysics Data System (ADS)
B?kowska, Karolina; Olech, Arkadiusz
2014-09-01
HT Cassiopeiae was discovered over seventy years ago (Hoffmeister 1943). Unfortunately, for 35 years this object did not receive any attention, until the eclipses of HT Cas were observed by Bond. After a first analysis, Patterson (1981) called HT Cas "a Rosetta stone among dwarf novae". Since then, the literature on this star is still growing, reaching several dozens of publications. We present an orbital period determination of HT Cas during the November 2010 super-outburst, but also during a longer time span, to check its stability.
Initial on-orbit radiometric calibration of the Suomi NPP VIIRS reflective solar bands
NASA Astrophysics Data System (ADS)
Lei, Ning; Wang, Zhipeng; Fulbright, Jon; Lee, Shihyan; McIntire, Jeff; Chiang, Kwofu; Xiong, Xiaoxiong
2012-09-01
The on-orbit radiometric response calibration of the VISible/Near InfraRed (VISNIR) and the Short-Wave InfraRed (SWIR) bands of the Visible/Infrared Imager/Radiometer Suite (VIIRS) aboard the Suomi National Polar-orbiting Partnership (NPP) satellite is carried out through a Solar Diffuser (SD). The transmittance of the SD screen and the SD's Bidirectional Reflectance Distribution Function (BRDF) are measured before launch and tabulated, allowing the VIIRS sensor aperture spectral radiance to be accurately determined. The radiometric response of a detector is described by a quadratic polynomial of the detector's digital number (dn). The coefficients were determined before launch. Once on orbit, the coefficients are assumed to change by a common factor: the F-factor. The radiance scattered from the SD allows the determination of the F-factor. In this Proceeding, we describe the methodology and the associated algorithms in the determination of the F-factors and discuss the results.
Precise Orbit Determination of the GOCE Re-Entry Phase
NASA Astrophysics Data System (ADS)
Gini, Francesco; Otten, Michiel; Springer, Tim; Enderle, Werner; Lemmens, Stijn; Flohrer, Tim
2015-03-01
During the last days of the GOCE mission, after the GOCE spacecraft ran out of fuel, it slowly decayed before finally re-entering the atmosphere on the 11th November 2013. As an integrated part of the AOCS, GOCE carried a GPS receiver that was in operations during the re-entry phase. This feature provided a unique opportunity for Precise Orbit Determination (POD) analysis. As part of the activities carried out by the Navigation Support Office (HSO-GN) at ESOC, precise ephemerides of the GOCE satellite have been reconstructed for the entire re-entry phase based on the available GPS observations of the onboard LAGRANGE receiver. All the data available from the moment the thruster was switched off on the 21st of October 2013 to the last available telemetry downlink on the 10th November 2013 have been processed, for a total of 21 daily arcs. For this period a dedicated processing sequence has been defined and implemented within the ESA/ESOC NAvigation Package for Earth Observation Satellites (NAPEOS) software. The computed results show a post-fit RMS of the GPS undifferenced carrier phase residuals (ionospheric-free linear combination) between 6 and 14 mm for the first 16 days which then progressively increases up to about 80 mm for the last available days. An orbit comparison with the Precise Science Orbits (PSO) generated at the Astronomical Institute of the University of Bern (AIUB, Bern, Switzerland) shows an average difference around 9 cm for the first 8 daily arcs and progressively increasing up to 17 cm for the following days. During this reentry phase (21st of October - 10th November 2013) a substantial drop in the GOCE altitude is observed, starting from about 230 km to 130 km where the last GPS measurements were taken. During this orbital decay an increment of a factor of 100 in the aerodynamic acceleration profile is observed. In order to limit the mis-modelling of the non-gravitational forces (radiation pressure and aerodynamic effects) the newly developed software ARPA (Aerodynamics and Radiation Pressure Analysis) has been adopted to compute the forces acting on GOCE. An overview of the software techniques and the results of its implementation is presented in this paper. The use of the ARPA modelling leads to an average reduction of the carrier phase post-fit RMS of about 2 mm and decrement of the difference with the PSO orbits of more than 1 cm.
Improved Space Object Orbit Determination Using CMOS Detectors
NASA Astrophysics Data System (ADS)
Schildknecht, T.; Peltonen, J.; SÃ¤nnti, T.; Silha, J.; Flohrer, T.
2014-09-01
CMOS-sensors, or in general Active Pixel Sensors (APS), are rapidly replacing CCDs in the consumer camera market. Due to significant technological advances during the past years these devices start to compete with CCDs also for demanding scientific imaging applications, in particular in the astronomy community. CMOS detectors offer a series of inherent advantages compared to CCDs, due to the structure of their basic pixel cells, which each contains their own amplifier and readout electronics. The most prominent advantages for space object observations are the extremely fast and flexible readout capabilities, feasibility for electronic shuttering and precise epoch registration, and the potential to perform image processing operations on-chip and in real-time. The major challenges and design drivers for ground-based and space-based optical observation strategies have been analyzed. CMOS detector characteristics were critically evaluated and compared with the established CCD technology, especially with respect to the above mentioned observations. Similarly, the desirable on-chip processing functionalities which would further enhance the object detection and image segmentation were identified. Finally, we simulated several observation scenarios for ground- and space-based sensor by assuming different observation and sensor properties. We will introduce the analyzed end-to-end simulations of the ground- and space-based strategies in order to investigate the orbit determination accuracy and its sensitivity which may result from different values for the frame-rate, pixel scale, astrometric and epoch registration accuracies. Two cases were simulated, a survey using a ground-based sensor to observe objects in LEO for surveillance applications, and a statistical survey with a space-based sensor orbiting in LEO observing small-size debris in LEO. The ground-based LEO survey uses a dynamical fence close to the Earth shadow a few hours after sunset. For the space-based scenario a sensor in a sun-synchronous LEO orbit, always pointing in the anti-sun direction to achieve optimum illumination conditions for small LEO debris, was simulated. For the space-based scenario the simulations showed a 20 130 % improvement of the accuracy of all orbital parameters when varying the frame rate from 1/3 fps, which is the fastest rate for a typical CCD detector, to 50 fps, which represents the highest rate of scientific CMOS cameras. Changing the epoch registration accuracy from a typical 20.0 ms for a mechanical shutter to 0.025 ms, the theoretical value for the electronic shutter of a CMOS camera, improved the orbit accuracy by 4 to 190 %. The ground-based scenario also benefit from the specific CMOS characteristics, but to a lesser extent.
First Attempt of Orbit Determination of SLR Satellites and Space Debris Using Genetic Algorithms
NASA Astrophysics Data System (ADS)
Deleflie, F.; Coulot, D.; Descosta, R.; Fernier, A.; Richard, P.
2013-08-01
We present an orbit determination method based on genetic algorithms. Contrary to usual estimation methods mainly based on least-squares methods, these algorithms do not require any a priori knowledge of the initial state vector to be estimated. These algorithms can be applied when a new satellite is launched or for uncatalogued objects that appear in images obtained from robotic telescopes such as the TAROT ones. We show in this paper preliminary results obtained from an SLR satellite, for which tracking data acquired by the ILRS network enable to build accurate orbital arcs at a few centimeter level, which can be used as a reference orbit ; in this case, the basic observations are made up of time series of ranges, obtained from various tracking stations. We show as well the results obtained from the observations acquired by the two TAROT telescopes on the Telecom-2D satellite operated by CNES ; in that case, the observations are made up of time series of azimuths and elevations, seen from the two TAROT telescopes. The method is carried out in several steps: (i) an analytical propagation of the equations of motion, (ii) an estimation kernel based on genetic algorithms, which follows the usual steps of such approaches: initialization and evolution of a selected population, so as to determine the best parameters. Each parameter to be estimated, namely each initial keplerian element, has to be searched among an interval that is preliminary chosen. The algorithm is supposed to converge towards an optimum over a reasonable computational time.
Initial Mars Orbiter Laser Altimeter (MOLA) Measurements of the Mars Surface and Atmosphere
NASA Technical Reports Server (NTRS)
Abshire, James B.; Sun, Xiaoli; Afzal, Robert S.
1998-01-01
The Mars Orbiter Laser Altimeter (MOLA) has made an initial set of measurements of the Mars surface and atmosphere. As of this writing 27 orbital passes have been completed, starting Sept. 15, 1997 on orbit Pass 3 and orbits 20-36 and beginning again on March 27, 1998 for orbit passes 203 - 212. The lidar is working well in Mars orbit, and its data show contiguous measurement profiles of the Mars surface to its maximum range of 786 km, an average pulse detection rate of > 99% under clear atmospheric conditions, and < 1 m range resolution. MOLA has profiled the shape and heights of a variety of interesting Mars surface features, including Olympus Mons, the flat northern plains of Mars, Valles Marineris and the northern polar ice cap. It has also detected and profiled a series of cloud layers which occur near the edge of the polar cap and near 60-70 deg N latitude. This is the first time clouds around another planet have been measured using lidar.
Filter parameter tuning analysis for operational orbit determination support
NASA Technical Reports Server (NTRS)
Dunham, J.; Cox, C.; Niklewski, D.; Mistretta, G.; Hart, R.
1994-01-01
The use of an extended Kalman filter (EKF) for operational orbit determination support is being considered by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). To support that investigation, analysis was performed to determine how an EKF can be tuned for operational support of a set of earth-orbiting spacecraft. The objectives of this analysis were to design and test a general purpose scheme for filter tuning, evaluate the solution accuracies, and develop practical methods to test the consistency of the EKF solutions in an operational environment. The filter was found to be easily tuned to produce estimates that were consistent, agreed with results from batch estimation, and compared well among the common parameters estimated for several spacecraft. The analysis indicates that there is not a sharply defined 'best' tunable parameter set, especially when considering only the position estimates over the data arc. The comparison of the EKF estimates for the user spacecraft showed that the filter is capable of high-accuracy results and can easily meet the current accuracy requirements for the spacecraft included in the investigation. The conclusion is that the EKF is a viable option for FDD operational support.
JASON-1 Precise Orbit Determination (POD)with SLR and DORIS Tracking
NASA Technical Reports Server (NTRS)
Zelensky, N. P.; Luthcke, S. B.; Rowlands, D. D.; Beckley, B. D.; Lemoine, Frank G.; Wang, Y. M.; Chinn, D. S.; Williams, T. A.
2002-01-01
Jason-1, the TOPEX/POSEIDON (T/P) radar altimeter follow-on, is intended to continue measurement of the ocean surface with the same, if not better accuracy. T/P has demonstrated that, the time variation of ocean topography can be determined with an accuracy of a few centimeters, thanks to the availability of highly accurate orbits based on SLR and DORIS tracking. For verification and cross-calibration, Jason-1, was initially injected into the T/P orbit, flying just 72 seconds ahead of T/P. This configuration lasted over 21 Jason cycles. In mid-August T/P was maneuvered into its final tandem configuration, a parallel groundtrack, in order to improve the combined coverage. Preliminary investigations using cycles 1-9, shown at the June 2002 SWT, indicated that nominal Jason orbits can achieve the 2-3 cm accuracy objective, however several puzzling aspects of SLR and DORIS measurement modeling were also observed. This paper presents recent analysis of Jason SLR+DORIS POD spanning more than 20 cycles, and revisits several of the more puzzling issues, including estimation of the Laser Retroreflector Array (LRA) offset. The accuracy of the orbits and of the measurement modeling are evaluated using several tests, including SLR, DORIS, and altimeter crossover residual analysis, altimeter collinear analysis, and direct comparison with GPS and other orbits. T/P POD results over the same period are used as a reference.
NASA Technical Reports Server (NTRS)
Kibler, J. F.; Green, R. N.; Young, G. R.; Kelly, M. G.
1974-01-01
A method has previously been developed to satisfy terminal rendezvous and intermediate timing constraints for planetary missions involving orbital operations. The method uses impulse factoring in which a two-impulse transfer is divided into three or four impulses which add one or two intermediate orbits. The periods of the intermediate orbits and the number of revolutions in each orbit are varied to satisfy timing constraints. Techniques are developed to retarget the orbital transfer in the presence of orbit-determination and maneuver-execution errors. Sample results indicate that the nominal transfer can be retargeted with little change in either the magnitude (Delta V) or location of the individual impulses. Additonally, the total Delta V required for the retargeted transfer is little different from that required for the nominal transfer. A digital computer program developed to implement the techniques is described.
NASA Astrophysics Data System (ADS)
Klimyk, Anatoliy; Patera, Jiri
2006-01-01
In the paper, properties of orbit functions are reviewed and further developed. Orbit functions on the Euclidean space En are symmetrized exponential functions. The symmetrization is fulfilled by a Weyl group corresponding to a Coxeter-Dynkin diagram. Properties of such functions will be described. An orbit function is the contribution to an irreducible character of a compact semisimple Lie group G of rank n from one of its Weyl group orbits. It is shown that values of orbit functions are repeated on copies of the fundamental domain F of the affine Weyl group (determined by the initial Weyl group) in the entire Euclidean space En. Orbit functions are solutions of the corresponding Laplace equation in En, satisfying the Neumann condition on the boundary of F. Orbit functions determine a symmetrized Fourier transform and a transform on a finite set of points.
NASA Technical Reports Server (NTRS)
Frauenholz, R. B.; Bhat, R. S.; Shapiro, B. E.; Leavitt, R. K.
1998-01-01
Since its' launch on August 10, 1992, the TOPEX/Poseidon satellite hs successfully observed the earth's ocean circulation using a combination of precision orbit determination (POD) and dual-frequency radar altimetry.
Advances in precision orbit determination of GRACE satellites
NASA Astrophysics Data System (ADS)
Bettadpur, Srinivas; Save, Himanshu; Kang, Zhigui
The twin Gravity Recovery And Climate Experiment (GRACE) satellites carry a complete suite of instrumentation essential for precision orbit determination (POD). Dense, continuous and global tracking is provided by the Global Positioning System receivers. The satellite orientation is measured using two star cameras. High precision measurements of non-gravitational accel-erations are provided by accelerometers. Satellite laser ranging (SLR) retroreflectors are used for collecting data for POD validation. Additional validation is provided by the highly precise K-Band ranging system measuring distance changes between the twin GRACE satellites. This paper presents the status of POD for GRACE satellites. The POD quality will be vali-dated using the SLR and K-Band ranging data. The POD quality improvement from upgraded modeling of the GPS observations, including the transition to the new IGS05 standards, will be discussed. In addition, the contributions from improvements in the gravity field modeling -partly arising out of GRACE science results -will be discussed. The aspects of these improve-ments that are applicable for the POD of other low-Earth orbiting satellites will be discussed as well.
NASA Astrophysics Data System (ADS)
Iasko, P. P.; Orlov, V. V.
2015-10-01
The region of initial conditions for close to periodic orbits is studied in the general three-body problem with components of equal mass and zero angular momentum. A method proposed earlier, based on minimization of a functional equal to the sum of the squares of the differences between the initial and current coordinates and velocities of the bodies, is used to search for such orbits. The search was conducted among orbits with periods T â‰¤ 2 000 Ï„, where Ï„ is the mean time for a component to cross the triple system. Elongated structures are found in the region of initial conditions, each of which corresponds to a certain periodic orbit. The detected structures seem to be conentrated along characteristic curves corresponding to the exact periodic orbits. A boundary zone of the initial conditions has been discovered, to the left and right of which orbits arising from the Schubart orbit and S orbit lie. Close to periodic orbits in the boundary zone possess the properties of both types of orbits. As a rule, these have periods of ~102 Ï„. Examples of trajectories of the bodies are presented. Dynamical and geometrical properties of the studied orbits are described.
Numerical comparison of Kalman filter algorithms - Orbit determination case study
NASA Technical Reports Server (NTRS)
Bierman, G. J.; Thornton, C. L.
1977-01-01
Numerical characteristics of various Kalman filter algorithms are illustrated with a realistic orbit determination study. The case study of this paper highlights the numerical deficiencies of the conventional and stabilized Kalman algorithms. Computational errors associated with these algorithms are found to be so large as to obscure important mismodeling effects and thus cause misleading estimates of filter accuracy. The positive result of this study is that the U-D covariance factorization algorithm has excellent numerical properties and is computationally efficient, having CPU costs that differ negligibly from the conventional Kalman costs. Accuracies of the U-D filter using single precision arithmetic consistently match the double precision reference results. Numerical stability of the U-D filter is further demonstrated by its insensitivity to variations in the a priori statistics.
NASA Technical Reports Server (NTRS)
Iona, Glenn; Butler, James; Guenther, Bruce; Graziani, Larissa; Johnson, Eric; Kennedy, Brian; Kent, Criag; Lambeck, Robert; Waluschka, Eugne; Xiong, Xiaoxiong
2012-01-01
A gradual, but persistent, decrease in the optical throughput was detected during the early commissioning phase for the Suomi National Polar-Orbiting Partnership (SNPP) Visible Infrared Imager Radiometer Suite (VIIRS) Near Infrared (NIR) bands. Its initial rate and unknown cause were coincidently coupled with a decrease in sensitivity in the same spectral wavelength of the Solar Diffuser Stability Monitor (SDSM) raising concerns about contamination or the possibility of a system-level satellite problem. An anomaly team was formed to investigate and provide recommendations before commissioning could resume. With few hard facts in hand, there was much speculation about possible causes and consequences of the degradation. Two different causes were determined as will be explained in this paper. This paper will describe the build and test history of VIIRS, why there were no indicators, even with hindsight, of an on-orbit problem, the appearance of the on-orbit anomaly, the initial work attempting to understand and determine the cause, the discovery of the root cause and what Test-As-You-Fly (TAYF) activities, can be done in the future to greatly reduce the likelihood of similar optical anomalies. These TAYF activities are captured in the lessons learned section of this paper.
Computer-based instruction and reference documentation system for the orbit determination program
NASA Technical Reports Server (NTRS)
Hintz, G. R.; Ryne, M.; Watkins, M.; Kenney, M.; Overoye, D.
2003-01-01
The Orbit Determination Program set has been used at the Jet Propulsion Laboratory for nearly half a century to enable precision navigation of interplanetary and earth-orbiting missions and to support a myriad of scientific investigations.
Orbit determination using single station SLR data assisted by telescope pointing data
NASA Astrophysics Data System (ADS)
Sun, MingGuo; Liu, ChengZhi; Li, Zhenwei; Liu, Yang
2011-11-01
Abstract: the orbit determination exclusively using single station SLR(Satellite Laser Ranging) data is unviable, which limits the application of SLR technology in the observation of space debris. The paper puts forward that the orbit determination can be achieved through using SLR data assisted by the proper weighted telescope pointing data of the SLR system. To process the SLR data of AJISAI satellite by the above method, the data consist of the orbital arcs with 3-day spans, the precision of orbit determination is about 10 cm. Discussion shows that space debris orbit determination by the above method is also feasible.
NASA Technical Reports Server (NTRS)
Lemoine, F. G.; Zelensky, N. P.; Luthcke, S. B.; Rowlands, D. D.; Beckley, B. D.; Klosko, S. M.
2006-01-01
Launched in the summer of 1992, TOPEX/POSEIDON (T/P) was a joint mission between NASA and the Centre National d Etudes Spatiales (CNES), the French Space Agency, to make precise radar altimeter measurements of the ocean surface. After the remarkably successful 13-years of mapping the ocean surface T/P lost its ability to maneuver and was de-commissioned January 2006. T/P revolutionized the study of the Earth s oceans by vastly exceeding pre-launch estimates of surface height accuracy recoverable from radar altimeter measurements. The precision orbit lies at the heart of the altimeter measurement providing the reference frame from which the radar altimeter measurements are made. The expected quality of orbit knowledge had limited the measurement accuracy expectations of past altimeter missions, and still remains a major component in the error budget of all altimeter missions. This paper describes critical improvements made to the T/P orbit time series over the 13-years of precise orbit determination (POD) provided by the GSFC Space Geodesy Laboratory. The POD improvements from the pre-launch T/P expectation of radial orbit accuracy and Mission requirement of 13-cm to an expected accuracy of about 1.5-cm with today s latest orbits will be discussed. The latest orbits with 1.5 cm RMS radial accuracy represent a significant improvement to the 2.0-cm accuracy orbits currently available on the T/P Geophysical Data Record (GDR) altimeter product.
NASA Astrophysics Data System (ADS)
Son, Ju Young; Jo, Jung Hyun; Choi, Jin; Kim, Bang-Yeop; Yoon, Joh-Na; Yim, Hong-Suh; Choi, Young-Jun; Park, Sun-Youp; Bae, Young Ho; Roh, Dong-Goo; Park, Jang-Hyun; Kim, Ji-Hye
2015-09-01
We estimated the orbit of the Communication, Ocean and Meteorological Satellite (COMS), a Geostationary Earth Orbit (GEO) satellite, through data from actual optical observations using telescopes at the Sobaeksan Optical Astronomy Observatory (SOAO) of the Korea Astronomy and Space Science Institute (KASI), Optical Wide field Patrol (OWL) at KASI, and the Chungbuk National University Observatory (CNUO) from August 1, 2014, to January 13, 2015. The astrometric data of the satellite were extracted from the World Coordinate System (WCS) in the obtained images, and geometrically distorted errors were corrected. To handle the optically observed data, corrections were made for the observation time, light-travel time delay, shutter speed delay, and aberration. For final product, the sequential filter within the Orbit Determination Tool Kit (ODTK) was used for orbit estimation based on the results of optical observation. In addition, a comparative analysis was conducted between the precise orbit from the ephemeris of the COMS maintained by the satellite operator and the results of orbit estimation using optical observation. The orbits estimated in simulation agree with those estimated with actual optical observation data. The error in the results using optical observation data decreased with increasing number of observatories. Our results are useful for optimizing observation data for orbit estimation.
Radar and Optical Sensor Data Fusion for Orbital Determination of HEO Objects
NASA Astrophysics Data System (ADS)
Fernandez Sanchez, J.; Aivar Garcia, L.; Agueda Mate, A.; Utzmann, J.; Bartsch, G.; Abreu, D.; Flohrer, T.
2013-08-01
The paper presents the results of a GSTP project led by GMV for ESA/ESOC to define and experimentally analyse orbit determination techniques for the cataloguing of objects in Highly Eccentric Orbits (HEO), such as the Geostationary Transfer Orbits (GTO) and Molniya-type orbits, using a combination, or fusion, of observations acquired by ground-based radars and optical telescopes. An experimental tracking campaign was scheduled and performed to test the evaluated concepts. Additionally, the needs of a future tracking network in terms of topology and sensors characteristics for the coverage of the population of HEO object were assessed and formulated. It is shown that acceptable orbit determination results for objects on eccentric orbits can only be expected when a longer arc of the orbit is covered with observations. As a result, the orbit determination of such objects would highly benefit from the combination of observations from optical telescopes and radars.
Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter
NASA Technical Reports Server (NTRS)
Slojkowski, Steven; Lowe, Jonathan; Woodburn, James
2015-01-01
Since launch, the FDF has performed daily OD for LRO using the Goddard Trajectory Determination System (GTDS). GTDS is a batch least-squares (BLS) estimator. The tracking data arc for OD is 36 hours. Current operational OD uses 200 x 200 lunar gravity, solid lunar tides, solar radiation pressure (SRP) using a spherical spacecraft area model, and point mass gravity for the Earth, Sun, and Jupiter. LRO tracking data consists of range and range-rate measurements from: Universal Space Network (USN) stations in Sweden, Germany, Australia, and Hawaii. A NASA antenna at White Sands, New Mexico (WS1S). NASA Deep Space Network (DSN) stations. DSN data was sparse and not included in this study. Tracking is predominantly (50) from WS1S. The OD accuracy requirements are: Definitive ephemeris accuracy of 500 meters total position root-mean-squared (RMS) and18 meters radial RMS. Predicted orbit accuracy less than 800 meters root sum squared (RSS) over an 84-hour prediction span.
Gravity and Tide Parameters Determined from Satellite and Spacecraft Orbits
NASA Astrophysics Data System (ADS)
Jacobson, Robert A.
2015-05-01
As part of our work on the development of the Jovian and Saturnian satellite ephemerides to support the Juno and Cassini missions, we determined a number of planetary system gravity parameters. This work did not take into account tidal forces. In fact, we saw no obvious observational evidence of tidal effects on the satellite or spacecraft orbits. However, Lainey et al. (2009 Nature 459, 957) and Lainey et. al (2012 Astrophys. J. 752, 14) have published investigations of tidal effects in the Jovian and Saturnian systems, respectively. Consequently, we have begun a re-examination of our ephemeris work that includes a model for tides raised on the planet by the satellites as well as tides raised on the satellites by the planet. In this paper we briefly review the observations used in our ephemeris production; they include astrometry from the late 1800s to 2014, mutual events, eclipses, occultatons, and data acquired by the Pioneer, Voyager, Ulysses, Cassini, Galileo, and New Horizons spacecraft. We summarize the gravity parameter values found from our original analyses. Next we discuss our tidal acceleration model and its impact on the gravity parameter determination. We conclude with preliminary results found when the reprocessing of the observations includes tidal forces acting on the satellites and spacecraft.
NASA Technical Reports Server (NTRS)
Forcey, W.; Minnie, C. R.; Defazio, R. L.
1995-01-01
The Geostationary Operational Environmental Satellite (GOES)-8 experienced a series of orbital perturbations from autonomous attitude control thrusting before perigee raising maneuvers. These perturbations influenced differential correction orbital state solutions determined by the Goddard Space Flight Center (GSFC) Goddard Trajectory Determination System (GTDS). The maneuvers induced significant variations in the converged state vector for solutions using increasingly longer tracking data spans. These solutions were used for planning perigee maneuvers as well as initial estimates for orbit solutions used to evaluate the effectiveness of the perigee raising maneuvers. This paper discusses models for the incorporation of attitude thrust effects into the orbit determination process. Results from definitive attitude solutions are modeled as impulsive thrusts in orbit determination solutions created for GOES-8 mission support. Due to the attitude orientation of GOES-8, analysis results are presented that attempt to absorb the effects of attitude thrusting by including a solution for the coefficient of reflectivity, C(R). Models to represent the attitude maneuvers are tested against orbit determination solutions generated during real-time support of the GOES-8 mission. The modeling techniques discussed in this investigation offer benefits to the remaining missions in the GOES NEXT series. Similar missions with large autonomous attitude control thrusting, such as the Solar and Heliospheric Observatory (SOHO) spacecraft and the INTELSAT series, may also benefit from these results.
Initial On-Orbit Radiometric Calibration of the Suomi NPP VIIRS Reflective Solar Bands
NASA Technical Reports Server (NTRS)
Lei, Ning; Wang, Zhipeng; Fulbright, Jon; Lee, Shihyan; McIntire, Jeff; Chiang, Vincent; Xiong, Jack
2012-01-01
The on-orbit radiometric response calibration of the VISible/Near InfraRed (VISNIR) and the Short-Wave InfraRed (SWIR) bands of the Visible/Infrared Imager/Radiometer Suite (VIIRS) aboard the Suomi National Polar-orbiting Partnership (NPP) satellite is carried out through a Solar Diffuser (SD). The transmittance of the SD screen and the SD's Bidirectional Reflectance Distribution Function (BRDF) are measured before launch and tabulated, allowing the VIIRS sensor aperture spectral radiance to be accurately determined. The radiometric response of a detector is described by a quadratic polynomial of the detector?s digital number (dn). The coefficients were determined before launch. Once on orbit, the coefficients are assumed to change by a common factor: the F-factor. The radiance scattered from the SD allows the determination of the F-factor. In this Proceeding, we describe the methodology and the associated algorithms in the determination of the F-factors and discuss the results.
Orbital Moment Determination in (MnxFe1-x)3O4 Nanoparticles
Pool, V. L.; Jolley, C.; Douglas, T.; Arenholz, E.; Idzerda, Y. U.
2010-10-22
Nanoparticles of (Mn{sub x}Fe{sub 1-x}){sub 3}O{sub 4} with a concentration ranging from x = 0 to 1 and a crystallite size of 14-15 nm were measured using X-ray absorption spectroscopy and X-ray magnetic circular dichroism to determine the ratio of the orbital moment to the spin moment for Mn and Fe. At low Mn concentrations, the Mn substitutes into the host Fe{sub 3}O{sub 4} spinel structure as Mn{sup 2+} in the tetrahedral A-site. The net Fe moment, as identified by the X-ray dichroism intensity, is found to increase at the lowest Mn concentrations then rapidly decrease until no dichroism is observed at 20% Mn. The average Fe orbit/spin moment ratio is determined to initially be negative and small for pure Fe{sub 3}O{sub 4} nanoparticles and quickly go to 0 by 5%-10% Mn addition. The average Mn moment is anti-aligned to the Fe moment with an orbit/spin moment ratio of 0.12 which gradually decreases with Mn concentration.
NASA Technical Reports Server (NTRS)
Rind, D.; Peteet, D.; Kukla, G.
1989-01-01
The possibility of initiating the growth of ice sheets by solar insolation variations is examined. The study is conducted using a climate model with three different orbital configurations corresponding to 116,000 and 106,000 yr before the present and a modified insolation field with greater reductions in summer insolation at high northern latitudes. Despite the reduced summer and fall insolation, the model fails to maintain snow cover through the summer at locations of suspected ice sheet initiation. The results suggest that there is a discrepancy between the model's response to Milankovitch perturbations and the geophysical evidence of ice sheet initiation. If the model results are correct, the growth of ice shown by geophysical evidence would have occurred in an extremely ablative environment, demanding a complicated strategy.
Dawn Orbit Determination Team: Trajectory Modeling and Reconstruction Processes at Vesta
NASA Technical Reports Server (NTRS)
Abrahamson, Matthew J.; Ardito, Alessandro; Han, Dongsuk; Haw, Robert; Kennedy, Brian; Mastrodemos, Nick; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The Dawn spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012. In order to maintain the designated science reference orbits and enable the transfers between those orbits, precise and timely orbit determination was required. Challenges included low-thrust ion propulsion modeling, estimation of relatively unknown Vesta gravity and rotation models, track-ing data limitations, incorporation of real-time telemetry into dynamics model updates, and rapid maneuver design cycles during transfers. This paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.
An Empirical State Error Covariance Matrix Orbit Determination Example
NASA Technical Reports Server (NTRS)
Frisbee, Joseph H., Jr.
2015-01-01
State estimation techniques serve effectively to provide mean state estimates. However, the state error covariance matrices provided as part of these techniques suffer from some degree of lack of confidence in their ability to adequately describe the uncertainty in the estimated states. A specific problem with the traditional form of state error covariance matrices is that they represent only a mapping of the assumed observation error characteristics into the state space. Any errors that arise from other sources (environment modeling, precision, etc.) are not directly represented in a traditional, theoretical state error covariance matrix. First, consider that an actual observation contains only measurement error and that an estimated observation contains all other errors, known and unknown. Then it follows that a measurement residual (the difference between expected and observed measurements) contains all errors for that measurement. Therefore, a direct and appropriate inclusion of the actual measurement residuals in the state error covariance matrix of the estimate will result in an empirical state error covariance matrix. This empirical state error covariance matrix will fully include all of the errors in the state estimate. The empirical error covariance matrix is determined from a literal reinterpretation of the equations involved in the weighted least squares estimation algorithm. It is a formally correct, empirical state error covariance matrix obtained through use of the average form of the weighted measurement residual variance performance index rather than the usual total weighted residual form. Based on its formulation, this matrix will contain the total uncertainty in the state estimate, regardless as to the source of the uncertainty and whether the source is anticipated or not. It is expected that the empirical error covariance matrix will give a better, statistical representation of the state error in poorly modeled systems or when sensor performance is suspect. In its most straight forward form, the technique only requires supplemental calculations to be added to existing batch estimation algorithms. In the current problem being studied a truth model making use of gravity with spherical, J2 and J4 terms plus a standard exponential type atmosphere with simple diurnal and random walk components is used. The ability of the empirical state error covariance matrix to account for errors is investigated under four scenarios during orbit estimation. These scenarios are: exact modeling under known measurement errors, exact modeling under corrupted measurement errors, inexact modeling under known measurement errors, and inexact modeling under corrupted measurement errors. For this problem a simple analog of a distributed space surveillance network is used. The sensors in this network make only range measurements and with simple normally distributed measurement errors. The sensors are assumed to have full horizon to horizon viewing at any azimuth. For definiteness, an orbit at the approximate altitude and inclination of the International Space Station is used for the study. The comparison analyses of the data involve only total vectors. No investigation of specific orbital elements is undertaken. The total vector analyses will look at the chisquare values of the error in the difference between the estimated state and the true modeled state using both the empirical and theoretical error covariance matrices for each of scenario.
TLE-Aided Orbit Determination Using Single-Station SLR Data
NASA Astrophysics Data System (ADS)
Liang, Zhi-peng; Liu, Cheng-zhi; Fan, Cun-bo; Sun, Ming-guo
2012-10-01
It is difficult to use the single-station satellite laser ranging (SLR) data for orbit determination, due to the singular geometrical distribution of the observations. The single-station data produced by performing the diffuse- reflection SLR on the earth-orbiting space debris are therefore ineffective for orbit improvement. To solve this problem, we propose an orbit determination method by using single-station SLR data in aid of the two-line element set (TLE). For verifying its feasibility, this method is implemented and applied to the orbit determination of the satellite Ajisai, using the single-station SLR data of five passes in one day and the corresponding TLE. And on this basis, the five-day orbit prediction is generated, the result indicates that the errors of predicted positions are less than 40 m. In addition, the potential application of this method in the orbit improvement of space debris is discussed.
12 CFR 404.8 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... grant or deny requests. The Freedom of Information and Privacy Office shall be responsible for search...: (i) The name, title, and signature of the person responsible for the determination; (ii)...
NASA Technical Reports Server (NTRS)
Marr, Greg C.
2003-01-01
Differencing multiple, simultaneous Tracking and Data Relay Satellite System (TDRSS) one-way Doppler passes can yield metric tracking data usable for orbit determination for (low-cost) spacecraft which do not have TDRSS transponders or local oscillators stable enough to allow the one-way TDRSS Doppler tracking data to be used for early mission orbit determination. Orbit determination error analysis results are provided for low Earth orbiting spacecraft for various early mission tracking scenarios.
Determining the Eccentricity of the Moon's Orbit without a Telescope
NASA Astrophysics Data System (ADS)
Krisciunas, Kevin
2010-01-01
Ancient Greek astronomers knew that Moon's distance from the Earth was not constant. Ptolemy's model of the Moon's motion implied that the Moon ranged in distance from 33 to 64 Earth radii. This implied that its angular size ranged nearly a factor of two. Tycho Brahe's model of the Moon's motion implied a smaller distance range, some Â±3 percent at syzygy. However, the ancient and Renaissance astronomers are notably silent on the subject of measuring the angular size of the Moon as a check on the implied range of distance from their models of the position of the Moon. Using a quarter-inch hole in a piece of cardboard that slides along a yardstick, we show that pre-telescopic astronomers could have measured an accurate mean value of the angular size of the Moon, and that they could have determined a reasonably accurate value of the eccentricity of the Moon's orbit. The principal calibration for each observer is to measure the apparent angular diameter of a 91 mm disk viewed at a distance of 10 meters, giving a true angular size of 31.3 arcmin (the Moon's mean angular size). Because the sighting hole is not much bigger than the size of one's pupil, each observer obtains a personal correction factor with which to scale the raw measures. If one takes data over the course of 7 lunations (7.5 anomalistic months), any systematic errors which are a function of phase should even out over the course of the observations. We find that the random error of an individual observation of Â±0.8 arcmin can be achieved.
Advanced stellar compass onboard autonomous orbit determination, preliminary performance.
Betto, Maurizio; JÃ¸rgensen, John L; JÃ¸rgensen, Peter S; Denver, Troelz
2004-05-01
Deep space exploration is in the agenda of the major space agencies worldwide; certainly the European Space Agency (SMART Program) and the American NASA (New Millennium Program) have set up programs to allow the development and the demonstration of technologies that can reduce the risks and the cost of deep space missions. From past experience, it appears that navigation is the Achilles heel of deep space missions. Performed on ground, this imposes considerable constraints on the entire system and limits operations. This makes it is very expensive to execute, especially when the mission lasts several years and, furthermore, it is not failure tolerant. Nevertheless, to date, ground navigation has been the only viable solution. The technology breakthrough of advanced star trackers, like the advanced stellar compass (ASC), might change this situation. Indeed, exploiting the capabilities of this instrument, the authors have devised a method to determine the orbit of a spacecraft autonomously, onboard, and without a priori knowledge of any kind. The solution is robust and fast. This paper presents the preliminary performance obtained during the ground testing in August 2002 at the Mauna Kea Observatories. The main goals were: (1) to assess the robustness of the method in solving autonomously, onboard, the position lost-in-space problem; (2) to assess the preliminary accuracy achievable with a single planet and a single observation; (3) to verify the autonomous navigation (AutoNav) module could be implemented into an ASC without degrading the attitude measurements; and (4) to identify the areas of development and consolidation. The results obtained are very encouraging. PMID:15220158
19 CFR 210.42 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-04-01
... section 337(j) of the Tariff Act. (2) On certain motions to declassify information. The decision of the administrative law judge granting a motion to declassify information, in whole or in part, shall be in the form... determination concerning a motion for temporary relief are governed by Â§Â§ 210.65 and 210.66. The disposition...
An Independent Orbit Determination Simulation for the OSIRIS-REx Asteroid Sample Return Mission
NASA Technical Reports Server (NTRS)
Getzandanner, Kenneth; Rowlands, David; Mazarico, Erwan; Antreasian, Peter; Jackman, Coralie; Moreau, Michael
2016-01-01
After arriving at the near-Earth asteroid (101955) Bennu in late 2018, the OSIRIS-REx spacecraft will execute a series of observation campaigns and orbit phases to accurately characterize Bennu and ultimately collect a sample of pristine regolith from its surface. While in the vicinity of Bennu, the OSIRIS-REx navigation team will rely on a combination of ground-based radiometric tracking data and optical navigation (OpNav) images to generate and deliver precision orbit determination products. Long before arrival at Bennu, the navigation team is performing multiple orbit determination simulations and thread tests to verify navigation performance and ensure interfaces between multiple software suites function properly. In this paper, we will summarize the results of an independent orbit determination simulation of the Orbit B phase of the mission performed to test the interface between the OpNav image processing and orbit determination software packages.
NASA Astrophysics Data System (ADS)
Sokova, I. A.; Sokov, E. N.; Roschina, E. A.; Rastegaev, D. A.; Kiselev, A. A.; Balega, Yu. Yu.; Gorshanov, D. L.; Malogolovets, E. V.; Dyachenko, V. V.; Maksimov, A. F.
2014-07-01
In this paper we present the orbital elements of Linus satellite of 22 Kalliope asteroid. Orbital element determination is based on the speckle interferometry data obtained with the 6-m BTA telescope operated by SAO RAS. We processed 9 accurate positions of Linus orbiting around the main component of 22 Kalliope between 10 and 16 December, 2011. In order to determine the orbital elements of the Linus we have applied the direct geometric method. The formal errors are about 5 mas. This accuracy makes it possible to study the variations of the Linus orbital elements influenced by different perturbations over the course of time. Estimates of six classical orbital elements, such as the semi-major axis of the Linus orbit a = 1109 ± 6 km, eccentricity e = 0.016 ± 0.004, inclination i = 101° ± 1° to the ecliptic plane and others, are presented in this work.
Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise
NASA Technical Reports Server (NTRS)
Pachura, Daniel A.; Vavrina, Matthew A.; Carpenter, J. R.; Wright, Cinnamon A.
2014-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the missions commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.
Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise
NASA Technical Reports Server (NTRS)
Pachura, Daniel A.; Vavrina, Matthew A.; Carpenter, J. Russell; Wright, Cinnamon A.
2014-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the mission's commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.
Orbit determination of the Comet Rendezvous/Asteroid Flyby mission - Post-rendezvous phases
NASA Technical Reports Server (NTRS)
Miller, James K.; Wood, Lincoln J.; Weeks, Connie J.
1989-01-01
Orbit determination during the post-rendezvous phases of the Comet Rendezvous/Asteroid Flyby mission is described. The orbit determination process is discussed, with emphasis placed on optical imaging of landmarks and Doppler tracking. Rotational dynamics are introduced for the cometary nucleus. State estimation errors are given for spacecraft trajectory prediction and cometary nucleus attitude prediction. Estimation errors are also given for parameters that describe the cometary nucleus such as moments of inertia and gravity harmonics. The orbit determination performance in support of science observations while in orbit about the nucleus is described.
Plasma Cell Neoplasm Manifesting Initially as a Sub-Cutaneous Supra-Orbital Swelling
Jaiswal, Riddhi; Agarwal, Garima; Singh, Sudhir
2016-01-01
Multiple myeloma is a plasma cell neoplasm seen usually in patients over 50 years of age. Some cases may be asymptomatic initially and are detected during a routine test like complete blood count. They only require a close follow-up and monitoring. However, around 1% of these monoclonal gammopathy of undetermined significance progress to multiple myeloma every year and then they need to be taken care of by chemotherapy, targeted therapy, bisphosphonates and 6 monthly urine and bone examinations. Here, we present a case of 35-year-old female with an initial symptom of a vague backache along with a left subcutaneous supra-orbital swelling which was diagnosed as multiple myeloma by aspiration cytology and confirmed by ancillary tests. She has since been on treatment with bortezomib and prednisone and is responding well.
Orbit determination and analysis of meteors recently observed by Finnish Fireball Network
NASA Astrophysics Data System (ADS)
Dmitriev, V.; Lupovla, V.; Gritsevich, M.; Lyytinen, E.; Mineeva, S.
2015-10-01
We perform orbit determination and analysis of three fireballs recently observed by Finnish Fireball Network (FFN). Precise orbit determination was performed by using integration of differential equations of motion. This technique was implemented into free distributable software "Meteor Toolkit". Accounting of several perturbing forces are discussed. Also estimation of accuracy of orbital elements was obtained by propagation of observational error with using covariance transformation. Long-term backward integration was provided as well.
Performance of OSC's initial Amtec generator design, and comparison with JPL's Europa Orbiter goals
Schock, A.; Noravian, H.; Or, C.; Kumar, V.
1998-07-01
The procedure for the analysis (with overpotential correction) of multitube AMTEC (Alkali Metal Thermal-to-Electrical Conversion) cells described in Paper IECEC 98-243 was applied to a wide range of multicell radioisotope space power systems. System design options consisting of one or two generators, each with 2, 3, or 4 stacked GPHS (General Purpose Heat Source) modules, identical to those used on previous NASA missions, were analyzed and performance-mapped. The initial generators analyzed by OSC had 8 AMTEC cells on each end of the heat source stack, with five beta-alumina solid electrolyte (BASE) tubes per cell. The heat source and converters in the Orbital generator designs are embedded in a thermal insulation system consisting of Min-K fibrous insulation surrounded by graded-length molybdenum multifoils. Detailed analyses in previous Orbital studies found that such an insulation system could reduce extraneous heat losses to about 10%. For the above design options, the present paper presents the system mass and performance (i.e., the EOM system efficiency and power output and the BOM evaporator and clad temperatures) for a wide range of heat inputs and load voltages, and compares the results with JPL's preliminary goals for the Europa Orbiter mission to be launched in November 2003. The analytical results showed that the initial 16-cell generator designs resulted in either excessive evaporator and clad temperatures and/or insufficient power outputs to meet the JPL-specified mission goals. The computed performance of modified OSC generators with different numbers of AMTEC cells, cell diameters, cell lengths, cell materials, BASE tube lengths, and number of tubes per cell are described in Paper IECEC.98.245 in these proceedings.
FIRST ORBIT AND MASS DETERMINATIONS FOR NINE VISUAL BINARIES
Ling, J. F.
2012-01-15
This paper presents the first published orbits and masses for nine visual double stars: WDS 00149-3209 (B 1024), WDS 01006+4719 (MAD 1), WDS 03130+4417 (STT 51), WDS 04357+3944 (HU 1084), WDS 19083+2706 (HO 98 AB), WDS 19222-0735 (A 102 AB), WDS 20524+2008 (HO 144), WDS 21051+0757 (HDS 3004 AB), and WDS 22202+2931 (BU 1216). Masses were calculated from the updated Hipparcos parallax data when available and sufficiently precise, or from dynamical parallaxes otherwise. Other physical and orbital properties are also discussed.
First Orbit and Mass Determinations for Nine Visual Binaries
NASA Astrophysics Data System (ADS)
Ling, J. F.
2012-01-01
This paper presents the first published orbits and masses for nine visual double stars: WDS 00149-3209 (B 1024), WDS 01006+4719 (MAD 1), WDS 03130+4417 (STT 51), WDS 04357+3944 (HU 1084), WDS 19083+2706 (HO 98 AB), WDS 19222-0735 (A 102 AB), WDS 20524+2008 (HO 144), WDS 21051+0757 (HDS 3004 AB), and WDS 22202+2931 (BU 1216). Masses were calculated from the updated Hipparcos parallax data when available and sufficiently precise, or from dynamical parallaxes otherwise. Other physical and orbital properties are also discussed.
Fractography: determining the sites of fracture initiation.
Mecholsky, J J
1995-03-01
Fractography is the analysis of fracture surfaces. Here, it refers to quantitative fracture surface analysis (FSA) in the context of applying the principles of fracture mechanics to the topography observed on the fracture surface of brittle materials. The application of FSA is based on the principle that encoded on the fracture surface of brittle materials is the entire history of the fracture process. It is our task to develop the skills and knowledge to decode this information. There are several motivating factors for applying our knowledge of FSA. The first and foremost is that there is specific, quantitative information to be obtained from the fracture surface. This information includes the identification of the size and location of the fracture initiating crack or defect, the stress state at failure, the existence, or not, of local or global residual stress, the existence, or not, of stress corrosion and a knowledge of local processing anomalies which affect the fracture process. The second motivating factor is that the information is free. Once a material is tested to failure, the encoded information becomes available. If we decide to observe the features produced during fracture then we are rewarded with much information. If we decide to ignore the fracture surface, then we are left to guess and/or reason as to the cause of the failure without the benefit of all of the possible information available. This paper addresses the application of quantitative fracture surface analysis to basic research, material and product development, and "trouble-shooting" of in-service failures. First, the basic principles involved will be presented. Next, the methodology necessary to apply the principles will be presented. Finally, a summary of the presentation will be made showing the applicability to design and reliability. PMID:8621031
Dawn Orbit Determination Team: Modeling and Fitting of Optical Data at Vesta
NASA Technical Reports Server (NTRS)
Kennedy, Brian; Abrahamson, Matt; Ardito, Alessandro; Haw, Robert; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The Dawn spacecraft was launched on September 27th, 2007. Its mission is to consecutively rendezvous with and observe the two largest bodies in the main asteroid belt, Vesta and Ceres. It has already completed over a year's worth of direct observations of Vesta (spanning from early 2011 through late 2012) and is currently on a cruise trajectory to Ceres, where it will begin scientific observations in mid-2015. Achieving this data collection required careful planning and execution from all Dawn operations teams. Dawn's Orbit Determination (OD) team was tasked with reconstruction of the as-flown trajectory as well as determination of the Vesta rotational rate, pole orientation and ephemeris, among other Vesta parameters. Improved knowledge of the Vesta pole orientation, specifically, was needed to target the final maneuvers that inserted Dawn into the first science orbit at Vesta. To solve for these parameters, the OD team used radiometric data from the Deep Space Network (DSN) along with optical data reduced from Dawn's Framing Camera (FC) images. This paper will de-scribe the initial determination of the Vesta ephemeris and pole using a combination of radiometric and optical data, and also the progress the OD team has made since then to further refine the knowledge of Vesta's body frame orientation and rate with these data.
Real-Time and Post-Processed Orbit Determination and Positioning
NASA Technical Reports Server (NTRS)
Bar-Sever, Yoaz E. (Inventor); Bertiger, William I. (Inventor); Dorsey, Angela R. (Inventor); Harvey, Nathaniel E. (Inventor); Lu, Wenwen (Inventor); Miller, Kevin J. (Inventor); Miller, Mark A. (Inventor); Romans, Larry J. (Inventor); Sibthorpe, Anthony J. (Inventor); Weiss, Jan P. (Inventor); Garcia Fernandez, Miquel (Inventor); Gross, Jason (Inventor)
2016-01-01
Novel methods and systems for the accurate and efficient processing of real-time and latent global navigation satellite systems (GNSS) data are described. Such methods and systems can perform orbit determination of GNSS satellites, orbit determination of satellites carrying GNSS receivers, positioning of GNSS receivers, and environmental monitoring with GNSS data.
NASA Technical Reports Server (NTRS)
Fuchs, A. J. (Editor)
1979-01-01
Onboard and real time image processing to enhance geometric correction of the data is discussed with application to autonomous navigation and attitude and orbit determination. Specific topics covered include: (1) LANDSAT landmark data; (2) star sensing and pattern recognition; (3) filtering algorithms for Global Positioning System; and (4) determining orbital elements for geostationary satellites.
Real-Time and Post-Processed Orbit Determination and Positioning
NASA Technical Reports Server (NTRS)
Bar-Sever, Yoaz E. (Inventor); Bertiger, William I. (Inventor); Dorsey, Angela R. (Inventor); Harvey, Nathaniel E. (Inventor); Lu, Wenwen (Inventor); Miller, Kevin J. (Inventor); Miller, Mark A. (Inventor); Romans, Larry J. (Inventor); Sibthorpe, Anthony J. (Inventor); Weiss, Jan P. (Inventor); Garcia Fernandez, Miquel (Inventor); Gross, Jason (Inventor)
2015-01-01
Novel methods and systems for the accurate and efficient processing of real-time and latent global navigation satellite systems (GNSS) data are described. Such methods and systems can perform orbit determination of GNSS satellites, orbit determination of satellites carrying GNSS receivers, positioning of GNSS receivers, and environmental monitoring with GNSS data.
Orbit determination of highly elliptical Earth orbiters using improved Doppler data-processing modes
NASA Technical Reports Server (NTRS)
Estefan, J. A.
1995-01-01
A navigation error covariance analysis of four highly elliptical Earth orbits is described, with apogee heights ranging from 20,000 to 76,800 km and perigee heights ranging from 1,000 to 5,000 km. This analysis differs from earlier studies in that improved navigation data-processing modes were used to reduce the radio metric data. For this study, X-band (8.4-GHz) Doppler data were assumed to be acquired from two Deep Space Network radio antennas and reconstructed orbit errors propagated over a single day. Doppler measurements were formulated as total-count phase measurements and compared to the traditional formulation of differenced-count frequency measurements. In addition, an enhanced data-filtering strategy was used, which treated the principal ground system calibration errors affecting the data as filter parameters. Results suggest that a 40- to 60-percent accuracy improvement may be achievable over traditional data-processing modes in reconstructed orbit errors, with a substantial reduction in reconstructed velocity errors at perigee. Historically, this has been a regime in which stringent navigation requirements have been difficult to meet by conventional methods.
Orbit Determination Error Analysis Results for the Triana Sun-Earth L2 Libration Point Mission
NASA Technical Reports Server (NTRS)
Marr, G.
2003-01-01
Using the NASA Goddard Space Flight Center's Orbit Determination Error Analysis System (ODEAS), orbit determination error analysis results are presented for all phases of the Triana Sun-Earth L1 libration point mission and for the science data collection phase of a future Sun-Earth L2 libration point mission. The Triana spacecraft was nominally to be released by the Space Shuttle in a low Earth orbit, and this analysis focuses on that scenario. From the release orbit a transfer trajectory insertion (TTI) maneuver performed using a solid stage would increase the velocity be approximately 3.1 km/sec sending Triana on a direct trajectory to its mission orbit. The Triana mission orbit is a Sun-Earth L1 Lissajous orbit with a Sun-Earth-vehicle (SEV) angle between 4.0 and 15.0 degrees, which would be achieved after a Lissajous orbit insertion (LOI) maneuver at approximately launch plus 6 months. Because Triana was to be launched by the Space Shuttle, TTI could potentially occur over a 16 orbit range from low Earth orbit. This analysis was performed assuming TTI was performed from a low Earth orbit with an inclination of 28.5 degrees and assuming support from a combination of three Deep Space Network (DSN) stations, Goldstone, Canberra, and Madrid and four commercial Universal Space Network (USN) stations, Alaska, Hawaii, Perth, and Santiago. These ground stations would provide coherent two-way range and range rate tracking data usable for orbit determination. Larger range and range rate errors were assumed for the USN stations. Nominally, DSN support would end at TTI+144 hours assuming there were no USN problems. Post-TTI coverage for a range of TTI longitudes for a given nominal trajectory case were analyzed. The orbit determination error analysis after the first correction maneuver would be generally applicable to any libration point mission utilizing a direct trajectory.
NASA Technical Reports Server (NTRS)
Doll, C.; Mistretta, G.; Hart, R.; Oza, D.; Cox, C.; Nemesure, M.; Bolvin, D.; Samii, Mina V.
1993-01-01
Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using the Goddard Trajectory Determination System (GTDS) and a real-time extended Kalman filter estimation system to process Tracking Data and Relay Satellite (TDRS) System (TDRSS) measurements in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. GTDS is the operational orbit determination system used by the FDD, and the extended Kalman fliter was implemented in an analysis prototype system, the Real-Time Orbit Determination System/Enhanced (RTOD/E). The Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generates an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the Geodynamics (GEODYN) orbit determination system with laser ranging tracking data. The TOPEX/Poseidon trajectories were estimated for the October 22 - November 1, 1992, timeframe, for which the latest preliminary POD results were available. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch cases were assessed using overlap comparisons, while the sequential cases were assessed with covariances and the first measurement residuals. The batch least-squares and forward-filtered RTOD/E orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 10 meters (m) for the batch least squares and less than 18 m for the sequential estimation solutions. The differences among the POD, GTDS, and RTOD/E solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.
Dawn Orbit Determination Team : Trajectory Modeling and Reconstruction Processes at Vesta
NASA Technical Reports Server (NTRS)
Abrahamson, Matt; Ardito, Alessandro; Han, Don; Haw, Robert; Kennedy, Brian; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The NASA Dawn spacecraft was launched on September 27, 2007 on a mission to study the asteroid belt's two largest objects, Vesta and Ceres. It is the first deep space orbiting mission to demonstrate solar-electric ion propulsion, providing the necessary delta-V to enable capture and escape from two extraterrestrial bodies. At this time, Dawn has completed its science campaign at Vesta and is currently on its journey to Ceres, where it will arrive in mid-2015. The spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012, capturing science data during four dedicated orbit phases. In order to maintain the reference orbits necessary for science and enable the transfers between those orbits, precise and timely orbit determination was required. The constraints associated with low-thrust ion propulsion coupled with the relatively unknown a priori gravity and rotation models for Vesta presented unique challenges for the Dawn orbit determination team. While [1] discusses the prediction performance of the orbit determination products, this paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Hodjatzadeh, M.; Samii, M. V.; Doll, C. E.; Hart, R. C.; Mistretta, G. D.
1991-01-01
The development of the Real-Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination on a Disk Operating System (DOS) based Personal Computer (PC) is addressed. The results of a study to compare the orbit determination accuracy of a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOD/E with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), is addressed. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for the Earth Radiation Budget Satellite (ERBS); the maximum solution differences were less than 25 m after the filter had reached steady state.
Applications of square-root information filtering and smoothing in spacecraft orbit determination
NASA Technical Reports Server (NTRS)
Wang, Tseng-Chan; Collier, James B.; Ekelund, John E.; Breckheimer, Peter J.
1988-01-01
The JPL (Jet Propulsion Laboratory) Orbit Determination Software System is a set of computer programs developed for the primary purpose of determining the flight path of deep-space mission spacecraft in NASA's Planetary Program and highly elliptical orbiting spacecraft in Earth orbit. The filtering processes available within the JPL Orbit Determination Software are discussed, and several examples are presented. In particular, solutions obtained by the Square Root Information Filter (SRIF) using Bierman's Estimation Subroutine Library (ESL) are discussed and compared with the solutions obtained by the singular value decomposition (SVD) technique. It is concluded that the SRIF filtering and smoothing algorithms are efficient and numerically stable for well-conditioned systems. The use of Bierman's ESL simplifies the task of maintaining the orbit determination software by providing efficient, tested filtering tools. For solving a large well-conditioned system (rank higher than 120), SRIF is approximately four times faster than SVD; however, for solving an ill-conditioned system, SVD is recommended.
OCO-2 (Orbiting Carbon Observatory-2) mission operations planning and initial operations experiences
NASA Astrophysics Data System (ADS)
Basilio, Ralph R.; Pollock, H. Randy; Hunyadi-Lay, Sarah L.
2014-10-01
OCO-2 (Orbiting Carbon Observatory-2) is the first NASA (National Aeronautics and Space Administration) mission dedicated to studying atmospheric carbon dioxide, specifically to identify sources (emitters) and sinks (absorbers) on a regional (1000 km x 1000 km) scale. The mission is designed to meet a science imperative by providing critical and urgent measurements needed to improve understanding of the carbon cycle and global climate change processes. The single instrument consisting of three grating spectrometers was built at the Jet Propulsion Laboratory, but is based on the design co-developed with Hamilton Sundstrand Corporation for the original OCO mission. The instrument underwent an extensive ground test program. This was generally made possible through the use of a thermal vacuum chamber with a window/port that allowed optical ground support equipment to stimulate the instrument. The instrument was later delivered to Orbital Sciences Corporation for integration and test with the LEOStar-2 spacecraft. During the overall ground test campaign, proper function and performance in simulated launch, ascent, and space environments were verified. The observatory was launched into space on 02 July 2014. Initial indications are that the instrument is meeting functional and performance specifications, and there is every expectation that the spatially-order, geo-located, calibrated spectra of reflected sunlight and the science retrievals will meet the Level 1 science requirements.
Determination of shuttle orbiter center of gravity from flight measurements
NASA Technical Reports Server (NTRS)
Hinson, E. W.; Nicholson, J. Y.; Blanchard, R. C.
1991-01-01
Flight measurements of pitch, yaw, and roll rates and the resultant rotationally induced linear accelerations during three orbital maneuvers on Shuttle mission space transportation system (STS) 61-C were used to calculate the actual orbiter center-of-gravity location. The calculation technique reduces error due to lack of absolute calibration of the accelerometer measurements and compensates for accelerometer temperature bias and for the effects of gravity gradient. Accuracy of the technique was found to be limited by the nonrandom and asymmetrical distribution of orbiter structural vibration at the accelerometer mounting location. Fourier analysis of the vibration was performed to obtain the power spectral density profiles which show magnitudes in excess of 10(exp 4) ug (sup 2)/Hz for the actual vibration and over 500 ug (sup 2)/Hz for the filtered accelerometer measurements. The data from this analysis provide a characterization of the Shuttle acceleration environment which may be useful in future studies related to accelerometer system application and zero-g investigations or processes.
GPS orbit determination at the National Geodetic Survey
NASA Technical Reports Server (NTRS)
Schenewerk, Mark S.
1992-01-01
The National Geodetic Survey (NGS) independently generates precise ephemerides for all available Global Positioning System (GPS) satellites. Beginning in 1991, these ephemerides were produced from double-differenced phase observations solely from the Cooperative International GPS Network (CIGNET) tracking sites. The double-difference technique combines simultaneous observations of two satellites from two ground stations effectively eliminating satellite and ground receiver clock errors, and the Selective Availability (S/A) signal degradation currently in effect. CIGNET is a global GPS tracking network whose primary purpose is to provide data for orbit production. The CIGNET data are collected daily at NGS and are available to the public. Each ephemeris covers a single week and is available within one month after the data were taken. Verification is by baseline repeatability and direct comparison with other ephemerides. Typically, an ephemeris is accurate at a few parts in 10(exp 7). This corresponds to a 10 meter error in the reported satellite positions. NGS is actively investigating methods to improve the accuracy of its orbits, the ultimate goal being one part in 10(exp 8) or better. The ephemerides are generally available to the public through the Coast Guard GPS Information Center or directly from NGS through the Geodetic Information Service. An overview of the techniques and software used in orbit generation will be given, the current status of CIGNET will be described, and a summary of the ephemeris verification results will be presented.
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2013 CFR
2013-04-01
... 18 Conservation of Power and Water Resources 2 2013-04-01 2012-04-01 true Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination. (a) Any individual whose request for a...
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2012 CFR
2012-04-01
... 18 Conservation of Power and Water Resources 2 2012-04-01 2012-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 18 Conservation of Power and Water Resources 2 2011-04-01 2011-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2014 CFR
2014-04-01
... 18 Conservation of Power and Water Resources 2 2014-04-01 2014-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2014 CFR
2014-04-01
... 20 Employees' Benefits 2 2014-04-01 2014-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2013 CFR
2013-04-01
... 20 Employees' Benefits 2 2013-04-01 2013-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2012 CFR
2012-04-01
... 20 Employees' Benefits 2 2012-04-01 2012-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 10 Energy 1 2010-01-01 2010-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations Â§ 9.29 Appeal from initial determination. (a) A requester may appeal a notice of denial of a Freedom of Information Act request for access to agency records,...
Orbit Determination of Non-cooperative Targets Using Laser Ranging Data at Changchun Station
NASA Astrophysics Data System (ADS)
Sun, J. N.; Liu, C. Z.; Fan, C. B.; Sun, M. G.
2015-09-01
The precise orbit determination software successfully processes the satellite laser ranging data of the non-cooperative targets in a single station. The insufficient observation data and the sole distribution data have become a principal difficulty in the orbit determination of the non-cooperative targets. Through the choices of dynamic models and the selections of solving parameters in the process of orbit determination, the condition equation can be solved with a convergence algorithm, and the orbit is obtained. The positional deviation obtained with the method of orbital overlap will be used as the accuracy index in calculating more groups of non-cooperative targets data. And the ranging deviation is obtained by comparing the trajectory information after orbit determination with the observation data uninvolved in orbit determination, which can be regarded as the externally coincident precision. The results show that the average ranging residual is 1.01 meters, the outer precision is 14.35 meters, and the precision of 1-day orbit prediction is 24.60 meters for non-cooperative target (4814).
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Feiertag, R.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1993-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite (TDRS) System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the May 18-24, 1992, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. During this period, there were two separate orbit-adjust maneuvers on one of the TDRSS spacecraft (TDRS-East) and one small orbit-adjust maneuver for Landsat-4. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 30 meters after the filter had reached steady state.
Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.
1991-01-01
The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.
Satellite orbit determination and gravity field recovery from satellite-to-satellite tracking
NASA Astrophysics Data System (ADS)
Wakker, K. F.; Ambrosius, B. A. C.; Leenman, H.
1989-07-01
Studies on satellite-to-satellite tracking (SST) with POPSAT (a geodetic satellite concept) and a ERS-class (Earth observation) satellite, a Satellite-to-Satellite Tracking (SST) gravity mission, and precise gravity field determination methods and mission requirements are reported. The first two studies primarily address the application of SST between the high altitude POPSAT and an ERS-class or GRM (Geopotential Research Mission) satellite to the orbit determination of the latter two satellites. Activities focussed on the determination of the tracking coverage of the lower altitude satellite by ground based tracking systems and by POPSAT, orbit determination error analysis and the determination of the surface forces acting on GRM. The third study surveys principles of SST, uncertainties of existing drag models, effects of direct luni-solar attraction and tides on orbit and the gravity gradient observable. Detailed ARISTOTELES (which replaced POPSAT) orbit determination error analyses were performed for various ground based tracking networks.
Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)
NASA Technical Reports Server (NTRS)
Mesarch, Michael A.; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James
2007-01-01
This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF's orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.
NASA Technical Reports Server (NTRS)
Wu, Jiun-Tsong; Yunck, Thomas P.
1992-01-01
A covariance analysis is presented for satellite tracking and gravity recovery with a differential Global Positioning System-based technique to be demonstrated on TOPEX in the early 1990s. The technique employs data from an ensemble of repeat ground tracks to recover a unique satellite epoch state for each track and a set of invariant positional parameters common to all tracks. The positional parameters represent the effect of mismodeled gravitational field on the satellite orbit. At an altitude of 1336 km, where gravity modeling is the dominant systematic error, averaging of random error over many arcs and adjustment of the gravity model reduce the final satellite position error. The positional parameters can then be used to produce a refined global gravity model. The analysis indicates that errors ranging from 5 to 8 cm in TOPEX altitude and 0.05 to 0.2 mGal for the gravity field can be achieved, depending on the number of repeat arcs used.
NASA Technical Reports Server (NTRS)
MacLeond, Todd C.; Sims, W. Herb; Varnavas,Kosta A.; Ho, Fat D.
2011-01-01
The Memory Test Experiment is a space test of a ferroelectric memory device on a low Earth orbit satellite that launched in November 2010. The memory device being tested is a commercial Ramtron Inc. 512K memory device. The circuit was designed into the satellite avionics and is not used to control the satellite. The test consists of writing and reading data with the ferroelectric based memory device. Any errors are detected and are stored on board the satellite. The data is sent to the ground through telemetry once a day. Analysis of the data can determine the kind of error that was found and will lead to a better understanding of the effects of space radiation on memory systems. The test is one of the first flight demonstrations of ferroelectric memory in a near polar orbit which allows testing in a varied radiation environment. The initial data from the test is presented. This paper details the goals and purpose of this experiment as well as the development process. The process for analyzing the data to gain the maximum understanding of the performance of the ferroelectric memory device is detailed.
NASA Technical Reports Server (NTRS)
Morinelli, Patrick; Cosgrove, jennifer; Blizzard, Mike; Nicholson, Ann; Robertson, Mika
2007-01-01
This paper provides an overview of the launch and early orbit activities performed by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of five probes comprising the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft. The FDF was tasked to support THEMIS in a limited capacity providing backup orbit determination support for validation purposes for all five THEMIS probes during launch plus 30 days in coordination with University of California Berkeley Flight Dynamics Center (UCB/FDC). The FDF's orbit determination responsibilities were originally planned to be as a backup to the UCB/FDC for validation purposes only. However, various challenges early on in the mission and a Spacecraft Emergency declared thirty hours after launch placed the FDF team in the role of providing the orbit solutions that enabled contact with each of the probes and the eventual termination of the Spacecraft Emergency. This paper details the challenges and various techniques used by the GSFC FDF team to successfully perform orbit determination for all five THEMIS probes during the early mission. In addition, actual THEMIS orbit determination results are presented spanning the launch and early orbit mission phase. Lastly, this paper enumerates lessons learned from the THEMIS mission, as well as demonstrates the broad range of resources and capabilities within the FDF for supporting critical launch and early orbit navigation activities, especially challenging for constellation missions.
NASA Technical Reports Server (NTRS)
Head, D. E.; Mitchell, K. L.
1967-01-01
Program computes the thermal environment of a spacecraft in a lunar orbit. The quantities determined include the incident flux /solar and lunar emitted radiation/, total radiation absorbed by a surface, and the resulting surface temperature as a function of time and orbital position.
Laser ranging network performance and routine orbit determination at D-PAF
NASA Technical Reports Server (NTRS)
Massmann, Franz-Heinrich; Reigber, C.; Li, H.; Koenig, Rolf; Raimondo, J. C.; Rajasenan, C.; Vei, M.
1993-01-01
ERS-1 is now about 8 months in orbit and has been tracked by the global laser network from the very beginning of the mission. The German processing and archiving facility for ERS-1 (D-PAF) is coordinating and supporting the network and performing the different routine orbit determination tasks. This paper presents details about the global network status, the communication to D-PAF and the tracking data and orbit processing system at D-PAF. The quality of the preliminary and precise orbits are shown and some problem areas are identified.
Analysis of orbit determination for space based optical space surveillance system
NASA Astrophysics Data System (ADS)
ScirÃ©, Gioacchino; Santoni, Fabio; Piergentili, Fabrizio
2015-08-01
The detection capability and orbit determination performance of a space based optical observation system exploiting the visible band is analyzed. The sensor characteristics, in terms of sensitivity and resolution are those typical of present state of the art star trackers. A mathematical model of the system has been built and the system performance assessed by numerical simulation. The selection of the observer satellite's has been done in order to maximize the number of observed objects in LEO, based on a statistical analysis of the space debris population in this region. The space objects' observability condition is analyzed and two batch estimator based on the Levenberg-Marquardt and on the Powell dog-leg algorithms have been implemented and their performance compared. Both the algorithms are sensitive to the initial guess. Its influence on the algorithms' convergence is assessed, showing that the Powell dog-leg, which is a trust region method, performs better.
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2014-12-01
Two methods that the author developed earlier for finding the intermediate perturbed orbit of a small celestial body from three pairs of range and range rate observations [1, 2] are applied to the determination of orbits of Near-Earth asteroids. The methods are based on using the superosculating orbits with three- and fourth-order tangency. The degrees of approximation of the real motion by the constructed intermediate orbits near the middle measurement time are two and three orders of magnitude higher than by the Keplerian orbit determined with the help of traditional methods. We calculated the orbits of the asteroids 99942 Apophis, 1566 Icarus, 4179 Toutatis, 2007 DN41 and 2012 DA14. For the sake of brevity, we call the method based on the orbit with third-order tangency as Algorithm A1 and the method based on the orbit with fourth-order tangency -- as Algorithm A2. The results of the calculations are compared with the results of the calculations by the version of the methods mentioned that allows us to construct the unperturbed Keplerian orbit. We call this version of the methods as Algorithm A. The observational data were simulated using the nominal trajectories of the selected asteroids. These trajectories were obtained by the numerical integration of the differential equations of motion subject to the perturbations from the eight major planets, Pluto, and the Moon. The integration was carried out with the help of the 15-order Everhart procedure [3]. The main results of the calculations are the following. When the reference time interval is shortened by half (for small sizes of this interval), the errors in the compared algorithms A, A1, A2 decrease approximately by the factors 4, 16, 64 in coordinates and by the factors 2, 8, 16 in velocities, respectively. Such behavior of the errors is most clearly seen with the asteroids 2007 DN41 and 2012 DA14. This leads to a significant increase in the accuracy of the real motion approximation by the intermediate orbits constructed using the A1 and A2 algorithms (2-4 orders of magnitude in coordinates and 4-7 orders of magnitude in velocities higher) compared to the accuracy of the approximation by Keplerian orbits with decreasing the reference arc of the trajectory. Here, the higher is the efficiency of the algorithms A1 and A2, the smaller are the values of the topocentric distances, i.e., the greater are the perturbations caused by the Earth's gravitation. The advantage of Algorithm A2 over Algorithm A1 in accuracy extends approximately one order of magnitude. The minimal methodic errors of the position vector by using the A1 and A2 algorithms range from several meters in the case of the asteroid Apophis to several millimeters in the case of the asteroid 2012 DA14. Hence, the numerical examples analyzed in this work lead us to conclude that the proposed in [1, 2] methods for determination of an intermediate perturbed orbit from range and range rate measurements at three time points allow for significantly raising the accuracy of the calculation of the initial asteroid orbits in comparison with the algorithm based on the finding the unperturbed Keplerian orbit. The shorter is the orbital arc specified by the extreme time points, the greater is the advantage of the algorithms suggested over the algorithms of the traditional approach in the accuracy. The advantage of the algorithms suggested in the accuracy increases with raising the perturbations too, which is especially important for calculation of the initial trajectories of the space objects detected in the Earth's neighbourhood. The work was supported by the Ministry of Education and Science of the Russian Federation, project no. 2014/223(1567).
Present status and future trends in near-Earth satellite orbit determination
NASA Technical Reports Server (NTRS)
Fuchs, A. J.
1981-01-01
The major components of an orbit determination system and the evolution of the elements making up each component are reviewed. Typical accuracies presently achievable in the orbit determination process, the factors limiting the accuracies, and improvements in the dynamic models used in the process are summarized. Models are developed for orbit determination programs which include: (1) time varying area for solar radiation pressure; (2) a time varying model for albedo radiation pressure; (3) Earth tides which account for the distortions in the Earth's body due to Sun and Moon attraction; and (4) ocean tides which affect satellite altimeter data.
Impact of tracking station distribution structure on BeiDou satellite orbit determination
NASA Astrophysics Data System (ADS)
Zhang, Rui; Zhang, Qin; Huang, Guanwen; Wang, Le; Qu, Wei
2015-11-01
The racking station distribution structure plays an important role in GNSS satellite orbit determination. Due to the current satellite distribution of the BeiDou satellite navigation system (BDS), the problem how to construct a reasonable distribution of tracking stations to obtain BDS satellite orbits with high precision has become a highly imperative issue. Based on the theory of dynamic orbit determination, two different station distributions were analyzed to study their impact on BDS precise and real-time orbit determination. Subsequently, the impact of Satellite Position Dilution of Precision (SPDOP) values on orbit determination was analyzed. Finally, an improved scheme for the tracking station distribution was designed based on the original scheme. The numerical results show that the SPDOP value can be used to evaluate the contribution of the tracking stations distribution on the BDS IGSO and MEO satellites orbit determination. In addition, the tracking stations which focus on the Asia-Pacific region play a key role in current BDS orbit determination.
Precise Orbit Determination for the GEOSAT Follow-On Spacecraft
NASA Technical Reports Server (NTRS)
Lemoine, Frank G.; Rowlands, David D.; Zelensky, Nikita P.; Luthcke, Scott B.; Cox, Christopher M.; Marr, Gregory C.
1999-01-01
The US Navy's GEOSAT Follow-On spacecraft was launched on February 10, 1998 with its primary mission objective to map the oceans using a radar altimeter. The spacecraft tracking complement consists of GPS receivers, a laser retroreflector and Doppler beacons. Since the GPS receivers have not yet returned reliable data, the only means of providing high-quality precise orbits has been though satellite laser ranging (SLR). SLR has tracked the spacecraft since April 22, 1998, and an average of 7 passes per day have been obtained from US and foreign stations. Since the predicted radial orbit error due to the gravity field is only two to three cm, the largest contributor to the high SLR residuals (10 cm) is the mismodelling of the non-conservative forces. The SLR residuals show a clear correlation with beta prime (solar elevation) angle, peaking in mid-August 1998 when the beta prime angle reached -80 to -90 degrees. We report in this paper on the analysis of the GFO tracking data (SLR, Doppler, and if available GPS) using GEODYN, and on the tuning of the non-conservative force model and the gravity model using these data.
Determining Mars parking orbits which ensure tangential periapsis burns at arrival and departure
NASA Technical Reports Server (NTRS)
Desai, Prasun N.; Buglia, James J.
1992-01-01
A method is presented which finds Mars parking orbits which allow tangential periapsis burns at both arrival and departure. This method accounts for the actual geometry at both arrival and departure between the hyperbolic asymptotes and the orbital plane, along with the precession effects caused by the oblateness of Mars. Thus, realistic Delta-V values (and hence initial low-earth orbit masses) are obtained for these orbits. The results obtained from the present method compare very well with a trajectory integration program while only requiring CPU times of about one minute. Therefore, due to the computational efficiency and accuracy, the present method would be an ideal tool to use in preliminary mission design, since it provides the opportunity to incorporate realistic Mars parking orbits effects.
NASA Technical Reports Server (NTRS)
Escher, William J. D.
1992-01-01
NASA's Earth-to-Orbit (ETO) Propulsion Technology Program, a multi-year/multi-task focused technology effort is, today, highly focused on conventional high-thrust cryogenic liquid chemical rocket engines and their envisioned future technology needs. But as highlighted in the U.S. National Ten-Year Space Launch Technology Plan, a set of less-conventional propulsion subjects, ones which offer significant promise for both, improving the state of the art and opening up new propulsion-capability possibilities, is now directed to the space propulsion planning community's attention. In conducting its forward-planning activities, it is highly appropriate that the ETO Program (and other programs as well) carefully consider integrating these "new initiative" subjects into the taskwork of future years. After an introductory consideration of the National Plan's propulsion-related directives, followed by a brief background overview of the ETO Program, the following specific new-initiative candidates are discussed from the standpoint of technology-program planning: operationally efficient propulsion systems; high-thrust hybrid rocket propulsion; low-cost, low-pressure expendable propulsion subsystems; advanced cryogenic in-space propulsion systems; integrated modular engine (IME) configured propulsion systems, and combined-cycle airbreathing/rocket propulsion systems.
Improved solution accuracy for TDRSS-based TOPEX/Poseidon orbit determination
NASA Technical Reports Server (NTRS)
Doll, C. E.; Mistretta, G. D.; Hart, R. C.; Oza, D. H.; Bolvin, D. T.; Cox, C. M.; Nemesure, M.; Niklewski, D. J.; Samii, M. V.
1994-01-01
Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using a batch-least-squares estimator available in the Goddard Trajectory Determination System (GTDS) and an extended Kalman filter estimation system to process Tracking and Data Relay Satellite (TDRS) System (TDRSS) measurements. GTDS is the operational orbit determination system used by the FDD in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. The extended Kalman filter was implemented in an orbit determination analysis prototype system, closely related to the Real-Time Orbit Determination System/Enhanced (RTOD/E) system. In addition, the Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generated an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the geodynamics (GEODYN) orbit determination system with laser ranging and Doppler Orbitography and Radiopositioning integrated by satellite (DORIS) tracking measurements. The TOPEX/Poseidon trajectories were estimated for November 7 through November 11, 1992, the timeframe under study. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch-least-squares solutions were assessed based on the solution residuals, while the sequential solutions were assessed based on primarily the estimated covariances. The batch-least-squares and sequential orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 2 meters for the batch-least-squares and less than 13 meters for the sequential estimation solutions. After the sequential estimation solutions were processed with a smoother algorithm, position differences with POD orbit solutions of less than 7 meters were obtained. The differences among the POD, GTDS, and filter/smoother solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.
Lunar Prospector Orbit Determination Uncertainties Using the High Resolution Lunar Gravity Models
NASA Technical Reports Server (NTRS)
Carranza, Eric; Konopliv, Alex; Ryne, Mark
1999-01-01
The Lunar Prospector (LP) mission began on January 6, 1998, when the LP spacecraft was launched from Cape Canaveral, Florida. The objectives of the mission were to determine whether water ice exists at the lunar poles, generate a global compositional map of the lunar surface, detect lunar outgassing, and improve knowledge of the lunar magnetic and gravity fields. Orbit determination of LP performed at the Jet Propulsion Laboratory (JPL) is conducted as part of the principal science investigation of the lunar gravity field. This paper will describe the JPL effort in support of the LP Gravity Investigation. This support includes high precision orbit determination, gravity model validation, and data editing. A description of the mission and its trajectory will be provided first, followed by a discussion of the orbit determination estimation procedure and models. Accuracies will be examined in terms of orbit-to-orbit solution differences, as a function of oblateness model truncation, and inclination in the plane-of-sky. Long term predictions for several gravity fields will be compared to the reconstructed orbits to demonstrate the accuracy of the orbit determination and oblateness fields developed by the Principal Gravity Investigator.
NASA Astrophysics Data System (ADS)
Waltner, Daniel; Gnutzmann, Sven; Tanner, Gregor; Richter, Klaus
2013-05-01
We study the implications of unitarity for pseudo-orbit expansions of the spectral determinants of quantum maps and quantum graphs. In particular, we advocate to group pseudo-orbits into subdeterminants. We show explicitly that the cancellation of long orbits is elegantly described on this level and that unitarity can be built in using a simple subdeterminant identity which has a nontrivial interpretation in terms of pseudo-orbits. This identity yields much more detailed relations between pseudo-orbits of different lengths than was known previously. We reformulate Newton identities and the spectral density in terms of subdeterminant expansions and point out the implications of the subdeterminant identity for these expressions. We analyze furthermore the effect of the identity on spectral correlation functions such as the autocorrelation and parametric cross-correlation functions of the spectral determinant and the spectral form factor.
NASA Astrophysics Data System (ADS)
Tupa, Peter R.; Quirin, S.; DeLeo, G. G.; McCluskey, G. E., Jr.
2007-12-01
We present a modified Fourier transform approach to determine the orbital parameters of detached visual binary stars. Originally inspired by Monet (ApJ 234, 275, 1979), this new method utilizes an iterative routine of refining higher order Fourier terms in a manner consistent with Keplerian motion. In most cases, this approach is not sensitive to the starting orbital parameters in the iterative loop. In many cases we have determined orbital elements even with small fragments of orbits and noisy data, although some systems show computational instabilities. The algorithm was constructed using the MAPLE mathematical software code and tested on artificially created orbits and many real binary systems, including Gliese 22 AC, Tau 51, and BU 738. This work was supported at Lehigh University by NSF-REU grant PHY-9820301.
NASA Technical Reports Server (NTRS)
Kennedy, Brian; Abrahamson, Matt; Ardito, Alessandro; Han, Dongsuk; Haw, Robert; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The Dawn spacecraft was launched on September 27th, 2007. Its mission is to consecutively rendezvous with and observe the two largest bodies in the asteroid belt, Vesta and Ceres. It has already completed over a year's worth of direct observations of Vesta (spanning from early 2011 through late 2012) and is currently on a cruise trajectory to Ceres, where it will begin scientific observations in mid-2015. Achieving this data collection required careful planning and execution from all spacecraft teams. Dawn's Orbit Determination (OD) team was tasked with accurately predicting the trajectory of the Dawn spacecraft during the Vesta science phases, and also determining the parameters of Vesta to support future science orbit design. The future orbits included the upcoming science phase orbits as well as the transfer orbits between science phases. In all, five science phases were executed at Vesta, and this paper will describe some of the OD team contributions to the planning and execution of those phases.
Orbiting Deep Space Relay Station (ODSRS). Volume 1: Requirement determination
NASA Technical Reports Server (NTRS)
Hunter, J. A.
1979-01-01
The deep space communications requirements of the post-1985 time frame are described and the orbiting deep space relay station (ODSRS) is presented as an option for meeting these requirements. Under current conditions, the ODSRS is not yet cost competitive with Earth based stations to increase DSN telemetry performance, but has significant advantages over a ground station, and these are sufficient to maintain it as a future option. These advantages include: the ability to track a spacecraft 24 hours per day with ground stations located only in the USA; the ability to operate at higher frequencies that would be attenuated by Earth's atmosphere; and the potential for building very large structures without the constraints of Earth's gravity.
A Neophyte's Determination of EY Ceph Curves and Orbital Constants
NASA Astrophysics Data System (ADS)
Menke, J. L.
2002-05-01
Using amateur level equipment, and with a beginner's knowledge of photometry, I have built the light curve for the eclipsing variable EY Cephei (subject suggested by Claud Lacy). This 10mag star has two 0.6mag eclipses per cycle, with a period of 7.97 days. The good observing news is that it is circumpolar (so easily visible), the bad news is that if the eclipses are not visible at night at your location (or if you miss them), you have to wait 6-8 months before they return. Combine this with only three days notice to start the campaign, no experience in precision photometry, and limited (but good) equipment, the amazing thing is that I achieved success in the first week (then the weather deteriorated...). I will present intensity curve results to date, discuss how I got to them, and provide the results of a preliminary calculation of the system parameters (orbits, sizes, etc).
A determination of the orbit of GX 301-2. [binary X-ray pulsars
NASA Technical Reports Server (NTRS)
Kelley, R.; Rappaport, S.; Petre, R.
1980-01-01
The pulse phase of GX 301-2(4U 1223-62) was tracked for 30 days with the SAS 3 satellite during 1979 January and February. It is suggested that most of the observed changes in pulse period are the result of Doppler shifts in a binary orbit, as opposed to changes in the intrinsic pulse period alone. The SAS 3 data allow orbital periods P(orb) equal to or greater than 23 days when a constant rate of change in the intrinsic pulse period is allowed as a free parameter in the orbital fits. For each trial orbital period the other orbital elements of the binary system are well determined. The SAS 3 data is combined with the Ariel 5 pulse arrival-time data to further restrict the allowed orbits. In both data sets a sharp minimum is observed in the Doppler delays of the pulse arrival times. Evidence is presented that the correct orbit is most likely the one with P(orb) = 35.0d, a projected semimajor axis for the neutron star of 304 light-seconds, and an eccentricity of 0.44. The relation of this system to the six X-ray binaries whose orbits have been determined previously is also discussed.
Improved solution accuracy for Landsat-4 (TDRSS-user) orbit determination
NASA Technical Reports Server (NTRS)
Oza, D. H.; Niklewski, D. J.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1994-01-01
This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using a Prototype Filter Smoother (PFS), with the accuracy of an established batch-least-squares system, the Goddard Trajectory Determination System (GTDS). The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and convariances for the sequential case) of solutions produced by the batch and sequential methods. The filtered and smoothed PFS orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 15 meters.
NASA Astrophysics Data System (ADS)
Peng, Dong-ju; Wu, Bin
2012-07-01
As a special approach to orbit determination for satellites with spaceborne GPS receivers, the kinematic Precise Orbit Determination (POD) is independent of any mechanical model (e.g., the Earth gravity field, atmospheric drag, solar radiation pressure, etc.), and thus especially suitable for the orbit determination of Low Earth Orbiting (LEO)satellites perturbed strongly bythe atmosphere. In this paper, based on the space-borne dual-frequency GPS data, we study the kinematic POD, discuss the pre-processing of the data, and construct an algorithm of zero-difference kinematic POD. Using the observational data from GRACE (Gravity Recovery And Climate Experiment) satellites covering the whole month of February 2008, we verify the effectiveness and reliability of this algorithm. The results show that the kinematic POD may attain an accuracy of about 5 cm (with respect to satellite laser ranging data), which is at the same level as the dynamic and reduced-dynamic PODs
Ren Shulin; Fu Yanning E-mail: fyn@pmo.ac.c
2010-05-15
Untill now, the Hipparcos intermediate astrometric data (HIAD) have contributed little to the full orbit determination of double-lined spectroscopic binaries (SB2s). This is because the photocenter of such a binary system is usually not far from the system mass center, and its orbital wobble is generally weak with respect to the accuracy of the HIAD. However, the HIAD have been recently revised and the accuracy is increased by a factor of 2.2 in the total weight. Therefore, it is interesting to see if the revised HIAD can be used in the orbit determination at least for some SB2s. In this paper, we first search the 9th Catalogue of Orbits of Spectroscopic Binaries (S{sub B{sup 9}}) for SB2s with reliable spectroscopic orbital solutions and with periods between 50 days and 3.2 years. This leaves us with 56 systems. The full orbital solutions of these systems are then determined from the HIAD by a highly efficient grid search method developed in this paper. The high efficiency is achieved by reducing the number of nonlinear model parameters to one, and by allowing all parameters to be adjustable within a region centered at each grid point. After a variety of tests, we finally accept orbital solutions of 13 systems. Among these systems, six (HIP 677, 20894, 87895, 95995, 101382, and 111170) are well resolved with reliable interferometric data. Orbital solutions from these data are consistent with our results. The full orbital solutions of the other seven systems (HIP 9121, 17732, 32040, 57029, 76006, 102431, and 116360) are determined for the first time.
TDRSS-user orbit determination using batch least-squares and sequential methods
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1993-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.
TDRSS-user orbit determination using batch least-squares and sequential methods
NASA Astrophysics Data System (ADS)
Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1993-02-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.
U.S. initiatives in the international effort to mitigate the orbital debris environment
NASA Astrophysics Data System (ADS)
Levin, George M.
1996-10-01
Following release of the 1989 'Report on Orbital Debris' by the Interagency Group (Space) for the National Security Council, NASA undertook a series of extensive bilateral discussions with the major spacefaring nations on the topic of orbital debris. These discussions led to a greater understanding of both the cause and the effect of orbital debris. As a result of these discussions, the major spacefaring nations have taken definitive steps to redesign their launch vehicles and spacecraft so as to mitigate the production of orbital debris. In 1993 the National Aeronautics and Space Administration (NASA), the European Space Agency (ESA), and Japan formed a multilateral Inter- Agency Orbital Debris Coordination Committee (IADC). Since that time the Russian Space Agency (RSA), the Chines National Space Agency (CNSA), the French National Space Agency (CNES), the British National Space Agency (BNSA), and the Indian Space Agency (ISRO) have jointed the IADC. In 1994 orbital debris discussions began in the United Nations under the auspices of the Scientific and Technical Subcommittee of the Committee on the Peaceful Uses of Outer Space (UNCOPUOS). In 1995 UNCOPUOS adopted a multi-year program for studying orbital debris. In 1993 the White House Office of Science and Technology Policy (OSTP) and the National Security Council (NSC) undertook an interagency review to revise and update the 1989 'Report on Orbital Debris.' In November 1995 Dr. John H. Gibbons, the Assistant to the President for Science and Technology, released the 'Interagency Report on Orbital Debris -- 1995.'
Radial orbit error reduction and sea surface topography determination using satellite altimetry
NASA Technical Reports Server (NTRS)
Engelis, Theodossios
1987-01-01
A method is presented in satellite altimetry that attempts to simultaneously determine the geoid and sea surface topography with minimum wavelengths of about 500 km and to reduce the radial orbit error caused by geopotential errors. The modeling of the radial orbit error is made using the linearized Lagrangian perturbation theory. Secular and second order effects are also included. After a rather extensive validation of the linearized equations, alternative expressions of the radial orbit error are derived. Numerical estimates for the radial orbit error and geoid undulation error are computed using the differences of two geopotential models as potential coefficient errors, for a SEASAT orbit. To provide statistical estimates of the radial distances and the geoid, a covariance propagation is made based on the full geopotential covariance. Accuracy estimates for the SEASAT orbits are given which agree quite well with already published results. Observation equations are develped using sea surface heights and crossover discrepancies as observables. A minimum variance solution with prior information provides estimates of parameters representing the sea surface topography and corrections to the gravity field that is used for the orbit generation. The simulation results show that the method can be used to effectively reduce the radial orbit error and recover the sea surface topography.
A demonstration of high precision GPS orbit determination for geodetic applications
NASA Technical Reports Server (NTRS)
Lichten, S. M.; Border, J. S.
1987-01-01
High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.
Precise orbit determination for NASA's earth observing system using GPS (Global Positioning System)
NASA Technical Reports Server (NTRS)
Williams, B. G.
1988-01-01
An application of a precision orbit determination technique for NASA's Earth Observing System (EOS) using the Global Positioning System (GPS) is described. This technique allows the geometric information from measurements of GPS carrier phase and P-code pseudo-range to be exploited while minimizing requirements for precision dynamical modeling. The method combines geometric and dynamic information to determine the spacecraft trajectory; the weight on the dynamic information is controlled by adjusting fictitious spacecraft accelerations in three dimensions which are treated as first order exponentially time correlated stochastic processes. By varying the time correlation and uncertainty of the stochastic accelerations, the technique can range from purely geometric to purely dynamic. Performance estimates for this technique as applied to the orbit geometry planned for the EOS platforms indicate that decimeter accuracies for EOS orbit position may be obtainable. The sensitivity of the predicted orbit uncertainties to model errors for station locations, nongravitational platform accelerations, and Earth gravity is also presented.
NASA Technical Reports Server (NTRS)
Gordon, Steven C.
1993-01-01
Spacecraft in orbit near libration point L1 in the Sun-Earth system are excellent platforms for research concerning solar effects on the terrestrial environment. One spacecraft mission launched in 1978 used an L1 orbit for nearly 4 years, and future L1 orbital missions are also being planned. Orbit determination and station-keeping are, however, required for these orbits. In particular, orbit determination error analysis may be used to compute the state uncertainty after a predetermined tracking period; the predicted state uncertainty levels then will impact the control costs computed in station-keeping simulations. Error sources, such as solar radiation pressure and planetary mass uncertainties, are also incorporated. For future missions, there may be some flexibility in the type and size of the spacecraft's nominal trajectory, but different orbits may produce varying error analysis and station-keeping results. The nominal path, for instance, can be (nearly) periodic or distinctly quasi-periodic. A periodic 'halo' orbit may be constructed to be significantly larger than a quasi-periodic 'Lissajous' path; both may meet mission requirements, but perhaps the required control costs for these orbits are probably different. Also for this spacecraft tracking and control simulation problem, experimental design methods can be used to determine the most significant uncertainties. That is, these methods can determine the error sources in the tracking and control problem that most impact the control cost (output); it also produces an equation that gives the approximate functional relationship between the error inputs and the output.
Improved treatment of GPS force parameters in precise orbit determination applications
NASA Technical Reports Server (NTRS)
Vigue, Yvonne; Lichten, Stephen M.; Muellerschoen, Ron J.; Blewitt, Geoff; Heflin, Michael B.
1993-01-01
Data collected from the International Global Positioning System (GPS) Service (IGS) have been processed at JPL to determine presise orbits for the satellites of the GPS. We have tested a filtering technique to improve modeling of GPS solar radiation pressure force parameters. The new approach improves orbit quality for eclipsing satellites by a factor of two, with typical results in the 25 - 50 cm range.
FEDS - An experiment with a microprocessor-based orbit determination system using TDRS data
NASA Technical Reports Server (NTRS)
Shank, D.; Pajerski, R.
1986-01-01
An experiment in microprocessor-based onboard orbit determination has been conducted at NASA's Goddard Space Flight Center. The experiment collected forward-link observation data in real time from a prototype transponder and performed orbit estimation on a typical low-earth scientific satellite. This paper discusses the hardware and organizational configurations of the experiment, the structure of the onboard software, the mathematical models, and the experiment results.
Early results from the TOPEX/POSEIDON GPS precise orbit determination demonstration
NASA Technical Reports Server (NTRS)
Bertiger, Willy; Wu, Sien; Yunck, Tom; Muellerschoen, Ron; Willis, Pascal; Bar-Sever, Yoaz; Davis, AB; Haines, Bruce; Munson, Tim; Lichten, Steve
1993-01-01
TOPEX/POSEIDON, a US/French oceanographic mission launched in August 1992, is the first earth satellite to carry a multi-channel, dual frequency Global Positioning System (GPS) receiver capable of making high precision P-code pseudorange and carrier phase measurements. The receiver was placed on TOPEX/POSEIDON as an experiment to demonstrate the potential of differential GPS tracking for subdecimeter orbit determination. In addition to the receiver, TOPEX/POSEIDON carries two flight-proven tracking systems to provide the operational precise orbit determination needed to meet the mission scientific requirements. These include a French-built one-way Doppler system known as DORIS (Doppler Orbitography and Radiopositioning Integrated by Satellite) and a circular ring of laser retroreflectors. Here we evaluate the quality of the GPS-determined orbits by examining post-fit residuals, orbit comparisons with DORIS, and orbit repeatability on overlapping data arcs. Overlapping data arcs with 6 hrs of common data out of a 30-hr arc have an average root-mean-square (RMS) altitude difference of 3.0 cm for 9 arcs. The average RMS altitude difference about the mean with a DORIS orbit was 5.7 cm.
Improving Fermi Orbit Determination and Prediction in an Uncertain Atmospheric Drag Environment
NASA Technical Reports Server (NTRS)
Vavrina, Matthew A.; Newman, Clark P.; Slojkowski, Steven E.; Carpenter, J. Russell
2014-01-01
Orbit determination and prediction of the Fermi Gamma-ray Space Telescope trajectory is strongly impacted by the unpredictability and variability of atmospheric density and the spacecraft's ballistic coefficient. Operationally, Global Positioning System point solutions are processed with an extended Kalman filter for orbit determination, and predictions are generated for conjunction assessment with secondary objects. When these predictions are compared to Joint Space Operations Center radar-based solutions, the close approach distance between the two predictions can greatly differ ahead of the conjunction. This work explores strategies for improving prediction accuracy and helps to explain the prediction disparities. Namely, a tuning analysis is performed to determine atmospheric drag modeling and filter parameters that can improve orbit determination as well as prediction accuracy. A 45% improvement in three-day prediction accuracy is realized by tuning the ballistic coefficient and atmospheric density stochastic models, measurement frequency, and other modeling and filter parameters.
Improving FermiI Orbit Determination and Prediction in an Uncertain Atmospheric Drag Environment
NASA Technical Reports Server (NTRS)
Vavrina, Matthew A.; Newman, Clark Patrick; Slojkowski, Steven E.; Carpenter, J. Russell
2014-01-01
Orbit determination and prediction of the Fermi Gamma-ray Space Telescope trajectory is strongly impacted by the unpredictability and variability of atmospheric density and the spacecrafts ballistic coefficient. Operationally, Global Positioning System point solutions are processed with an extended Kalman filter for orbit determination, and predictions are generated for conjunction assessment with secondary objects. When these predictions are compared to Joint Space Operations Center radar-based solutions, the close approach distance between the two predictions can greatly differ ahead of the conjunction. This work explores strategies for improving prediction accuracy and helps to explain the prediction disparities. Namely, a tuning analysis is performed to determine atmospheric drag modeling and filter parameters that can improve orbit determination as well as prediction accuracy. A 45 improvement in three-day prediction accuracy is realized by tuning the ballistic coefficient and atmospheric density stochastic models, measurement frequency, and other modeling and filter parameters.
Orbit determination accuracies using satellite-to-satellite tracking
NASA Technical Reports Server (NTRS)
Vonbun, F. O.; Argentiero, P. D.; Schmid, P. E.
1977-01-01
The uncertainty in relay satellite sate is a significant error source which cannot be ignored in the reduction of satellite-to-satellite tracking data. Based on simulations and real data reductions, it is numerically impractical to use simultaneous unconstrained solutions to determine both relay and user satellite epoch states. A Bayesian or least squares estimation technique with an a priori procedure is presented which permits the adjustment of relay satellite epoch state in the reduction of satellite-to-satellite tracking data without the numerical difficulties introduced by an ill-conditioned normal matrix.
A review of GPS-based tracking techniques for TDRS orbit determination
NASA Technical Reports Server (NTRS)
Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S.-C.
1993-01-01
This article evaluates two fundamentally different approaches to the Tracking and Data Relay Satellite (TDRS) orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRS. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRS's broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and Tracking and Data Relay Satellite System satellites by ground receivers. Both strategies can be designed to meet future operational requirements for TDRS-II orbit determination.
GOCE orbit analysis: Long-wavelength gravity field determination using the acceleration approach
NASA Astrophysics Data System (ADS)
Baur, O.; Reubelt, T.; Weigelt, M.; Roth, M.; Sneeuw, N.
2012-08-01
The restricted sensitivity of the Gravity field and steady-state Ocean Circulation Explorer (GOCE) gradiometer instrument requires satellite gravity gradiometry to be supplemented by orbit analysis in order to resolve long-wavelength features of the geopotential. For the hitherto published releases of the GOCE time-wise (TIM) and GOCE space-wise gravity field seriesâ€”two of the official ESA productsâ€”the energy conservation method has been adopted to exploit GPS-based satellite-to-satellite tracking information. On the other hand, gravity field recovery from data collected by the CHAllenging Mini-satellite Payload (CHAMP) satellite showed the energy conservation principle to be a sub-optimal choice. For this reason, we propose to estimate the low-frequency part of the gravity field by the point-wise solution of Newton's equation of motion, also known as the acceleration approach. This approach balances the gravitational vector with satellite accelerations, and hence is characterized by (second-order) numerical differentiation of the kinematic orbit. In order to apply the method to GOCE, we present tailored processing strategies with regard to low-pass filtering, variance-covariance information handling, and robust parameter estimation. By comparison of our GIWF solutions (initials GI for "GeodÃ¤tisches Institut" and IWF for "Institut fÃ¼r WeltraumForschung") and the GOCE-TIM estimates with a state-of-the-art gravity field solution derived from GRACE (Gravity Recovery And Climate Experiment), we conclude that the acceleration approach is better suited for GOCE-only gravity field determination as opposed to the energy conservation method.
Experimental Study on the Precise Orbit Determination of the BeiDou Navigation Satellite System
He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald
2013-01-01
The regional service of the Chinese BeiDou satellite navigation system is now in operation with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Besides the standard positioning service with positioning accuracy of about 10 m, both precise relative positioning and precise point positioning are already demonstrated. As is well known, precise orbit and clock determination is essential in enhancing precise positioning services. To improve the satellite orbits of the BeiDou regional system, we concentrate on the impact of the tracking geometry and the involvement of MEOs, and on the effect of integer ambiguity resolution as well. About seven weeks of data collected at the BeiDou Experimental Test Service (BETS) network is employed in this experimental study. Several tracking scenarios are defined, various processing schemata are designed and carried out; and then, the estimates are compared and analyzed in detail. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. The involvement of MEOs and ambiguity-fixing also make the orbits better. PMID:23529116
Experimental study on the precise orbit determination of the BeiDou navigation satellite system.
He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald
2013-01-01
The regional service of the Chinese BeiDou satellite navigation system is now in operation with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Besides the standard positioning service with positioning accuracy of about 10 m, both precise relative positioning and precise point positioning are already demonstrated. As is well known, precise orbit and clock determination is essential in enhancing precise positioning services. To improve the satellite orbits of the BeiDou regional system, we concentrate on the impact of the tracking geometry and the involvement of MEOs, and on the effect of integer ambiguity resolution as well. About seven weeks of data collected at the BeiDou Experimental Test Service (BETS) network is employed in this experimental study. Several tracking scenarios are defined, various processing schemata are designed and carried out; and then, the estimates are compared and analyzed in detail. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. The involvement of MEOs and ambiguity-fixing also make the orbits better. PMID:23529116
NASA Technical Reports Server (NTRS)
Mashiku, Alinda; Garrison, James L.; Carpenter, J. Russell
2012-01-01
The tracking of space objects requires frequent and accurate monitoring for collision avoidance. As even collision events with very low probability are important, accurate prediction of collisions require the representation of the full probability density function (PDF) of the random orbit state. Through representing the full PDF of the orbit state for orbit maintenance and collision avoidance, we can take advantage of the statistical information present in the heavy tailed distributions, more accurately representing the orbit states with low probability. The classical methods of orbit determination (i.e. Kalman Filter and its derivatives) provide state estimates based on only the second moments of the state and measurement errors that are captured by assuming a Gaussian distribution. Although the measurement errors can be accurately assumed to have a Gaussian distribution, errors with a non-Gaussian distribution could arise during propagation between observations. Moreover, unmodeled dynamics in the orbit model could introduce non-Gaussian errors into the process noise. A Particle Filter (PF) is proposed as a nonlinear filtering technique that is capable of propagating and estimating a more complete representation of the state distribution as an accurate approximation of a full PDF. The PF uses Monte Carlo runs to generate particles that approximate the full PDF representation. The PF is applied in the estimation and propagation of a highly eccentric orbit and the results are compared to the Extended Kalman Filter and Splitting Gaussian Mixture algorithms to demonstrate its proficiency.
Phase Function Determination in Support of Orbital Debris Size Estimation
NASA Technical Reports Server (NTRS)
Hejduk, M. D.; Cowardin, H. M.; Stansbery, Eugene G.
2012-01-01
To recover the size of a space debris object from photometric measurements, it is necessary to determine its albedo and basic shape: if the albedo is known, the reflective area can be calculated; and if the shape is known, the shape and area taken together can be used to estimate a characteristic dimension. Albedo is typically determined by inferring the object s material type from filter photometry or spectroscopy and is not the subject of the present study. Object shape, on the other hand, can be revealed from a time-history of the object s brightness response. The most data-rich presentation is a continuous light-curve that records the object s brightness for an entire sensor pass, which could last for tens of minutes to several hours: from this one can see both short-term periodic behavior as well as brightness variations with phase angle. Light-curve interpretation, however, is more art than science and does not lend itself easily to automation; and the collection method, which requires single-object telescope dedication for long periods of time, is not well suited to debris survey conditions. So one is led to investigate how easily an object s brightness phase function, which can be constructed from the more survey-friendly point photometry, can be used to recover object shape. Such a recovery is usually attempted by comparing a phase-function curve constructed from an object s empirical brightness measurements to analytically-derived curves for basic shapes or shape combinations. There are two ways to accomplish this: a simple averaged brightness-versus phase curve assembled from the empirical data, or a more elaborate approach in which one is essentially calculating a brightness PDF for each phase angle bin (a technique explored in unpublished AFRL/RV research and in Ojakangas 2011); in each case the empirical curve is compared to analytical results for shapes of interest. The latter technique promises more discrimination power but requires more data; the former can be assembled in its essentials from fewer measurements but will be less definitive in its assignments. The goal of the present study is to evaluate both techniques under debris survey conditions to determine their relative performance and, additionally, to learn precisely how a survey should be conducted in order to maximize their performance. Because the distendedness of objects has more of an effect than their precise shape in calculating a characteristic dimension, one is interested in the techniques discrimination ability to distinguish between an elongated rectangular prism and a short rectangular prism or cube, or an elongated cylinder from a squat cylinder or sphere. Sensitivity studies using simulated data will be conducted to determine discrimination power for both techniques as a function of amount of data collected and range (and specific region) of phase angles sampled. Empirical GEODSS photometry data for distended objects (dead payloads with solar panels, rocket bodies) and compact objects (cubesats, calibration spheres, squat payloads) will also be used to test this discrimination ability. The result will be a recommended technique and data collection paradigm for debris surveys in order to maximize this type of discrimination.
Goetz, A.F.H.; Rowan, L.C.; Kingston, M.J.
1982-01-01
A shuttle-borne radiometer containing ten channels in the reflective infrared has demonstrated that direct identification of carbonates and hydroxyl-bearing minerals is possible by remote measurement from Earth orbit. Copyright ?? 1982 AAAS.
GOCE Precise Orbit Determination for the Entire Mission- Challenges in the Final Mission Phase
NASA Astrophysics Data System (ADS)
Jaggi, A.; Bock, H.; Meyer, U.
2015-03-01
The Gravity field and steady-state Ocean Circulation Explorer (GOCE), ESAâ€™s first Earth Explorer core mission, was launched on March 17, 2009 into a sun-synchronous dusk-dawn orbit and eventually re-entered into the Earthâ€™s atmosphere on November 11, 2013. A precise science orbit (PSO) product was provided by the GOCE High-level Processing Facility (HPF) from the GPS high-low Satellite-to-Satellite Tracking (hl-SST) data from the beginning until the very last days of the mission. We recapitulate the PSO procedure and refer to the results achieved until the official end of the GOCE mission on October 21, 2013, where independent validations with Satellite Laser Ranging (SLR) measurements confirmed a high quality of the PSO product of about 2 cm 1-D RMS. We then focus on the period after the official end of the mission, where orbits could still be determined thanks to the continuously running GPS receivers delivering high quality data until a few hours before the re-entry into the Earthâ€™s atmosphere. We address the challenges encountered for orbit determination during these last days and report on adaptions in the PSO procedure to also obtain good orbit results at the unprecedented low orbital altitudes below 224 km.
NASA Technical Reports Server (NTRS)
Peters, Palmer N.; Gregory, John C.
1992-01-01
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.
NASA Astrophysics Data System (ADS)
Cretaux, J.-F.; Nouel, F.; Valorge, C.; Janniere, P.
1994-05-01
The theory of perturbations suggests that, in the calculation of ephemerides, most errors due to mismodeling of the forces acting on a spacecraft are of a resonant nature. Colombo (1986; 1989) has shown that they can be corrected by adjusting a certain number of parameters relative to a simple empirical force inferred from the so-called Hill's equations in spite of the complexity of the error causes: mismodeling of the gravitational field, radiation pressure etc. This principle can not be extended to all types of orbits and are valid only for circular ones (ex: geostationary or low Earth orbit). This force was introduced into an orbit determination software and it was tested on the orbits of the LAGEOS, STARLETTE, SPOT2, TOPEX and finally GPS satellites.
Interplanetary Departure Stage Navigation by Means of Liaison Orbit Determination Architecture
NASA Technical Reports Server (NTRS)
McGranaghan, Ryan M.; Leonard, Jason M.; Fujimoto, Kohei; Parker, Jeffrey S.; Anderson, Rodney L.; Born, George H.
2013-01-01
Autonomous orbit determination for departure stages of interplanetary trajectories is conducted by means of realistic radiometric observations between the departing spacecraft and a satellite orbiting the first lunar libration point. Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) is used to estimate the orbit solution. This paper uses high-fidelity simulations to explore the utilization of LiAISON in providing improved accuracy for interplanetary departure missions. The use of autonomous navigation to supplement current techniques for interplanetary spacecraft is assessed using comparisons with groundbased navigation. Results from simulations including the Mars Science Laboratory, Mars Exploration Rover, and Cassini are presented. It is shown that observations from a dedicated LiAISON navigation satellite could be used to supplement ground-based measurements and significantly improve tracking performance.
DETERMINATION OF ORBITAL ELEMENTS OF SPECTROSCOPIC BINARIES USING HIGH-DISPERSION SPECTROSCOPY
Katoh, Noriyuki; Itoh, Yoichi; Toyota, Eri; Sato, Bun'ei
2013-02-01
Orbital elements of 37 single-lined spectroscopic binary systems (SB1s) and 5 double-lined spectroscopic binary systems (SB2s) were determined using high-dispersion spectroscopy. To determine the orbital elements accurately, we carried out precise Doppler shift measurements using the HIgh Dispersion Echelle Spectrograph mounted on the Okayama Astrophysical Observatory 1.88 m telescope. We achieved a radial-velocity precision of {approx}10 m s{sup -1} over seven years of observations. The targeted binaries have spectral types between F5 and K3, and are brighter than the 7th magnitude in the V band. The orbital elements of 28 SB1s and 5 SB2s were determined at least 10 times more precisely than previous measurements. Among the remaining nine SB1s, five objects were found to be single stars, and the orbital elements of four objects were not determined because our observations did not cover the entire orbital period. We checked the absorption lines from the secondary star for 28 SB1s and found that three objects were in fact SB2s.
Cassini Orbit Determination Performance during Saturn Satellite Tour: August 2005 - January 2006
NASA Technical Reports Server (NTRS)
Antreasian, Peter G.; Bordi, J. J.; Criddle, K. E.; Ionasescu, R.; Jacobson, R. A.; Jones, J. B.; MacKenzie, R. A.; Parcher, D. W.; Pelletier, F. J.; Roth, D. C.; Stauch, J. R.
2007-01-01
During the period spanning the second Enceladus flyby in July 2005 through the eleventh Titan encounter in January 2006, the Cassini spacecraft was successfully navigated through eight close-targeted satellite encounters. Three of these encounters included the 500 km flybys of the icy satellites Hyperion, Dione and Rhea and five targeted flybys of Saturn's largest moon, Titan. This paper will show how our refinements to Saturn's satellite ephemerides have improved orbit determination predictions. These refinements include the mass estimates of Saturn and its satellites by better than 0.5%. Also, it will be shown how this better orbit determination performance has helped to eliminate several statistical maneuvers that were scheduled to clean-up orbit determination and/or maneuver-execution errors.
Radio metric orbit determination for the Giotto mission to Comet Halley
NASA Technical Reports Server (NTRS)
Wood, L. J.; Mottinger, N. A.; Jordan, J. F.
1983-01-01
An international fleet of five spacecraft will fly past Comet Halley as it travels through the inner solar system in early 1986. This paper discusses orbit determination problems associated with the Giotto spacecraft, sponsored by the European Space Agency. The large number of spin axis precession maneuvers required to maintain the desired spacecraft attitude creates a new kind of radio metric orbit determination problem for this mission. This paper investigates the accuracy with which the Giotto spacecraft orbit can be determined relative to the earth or the sun, and establishes the sensitivity of this accuracy to the selection of the parameters to be estimated, the form of estimator used, the number of tracking stations employed, the length of the data arc, the selection of data types processed, and the levels of various error sources.
Evaluation the initial estimators using deterministic minimum covariance determinant algorithm
NASA Astrophysics Data System (ADS)
Alrawashdeh, Mufda Jameel; Sabri, Shamsul Rijal Muhammad; Ismail, Mohd Tahir
2014-07-01
The aim of the study is to examine five initial estimators introduced by Hubert et al. [1] with five additional new initial estimators by using the Deterministic Minimum Covariance Determinant algorithm, DetMCD. The objective of the DetMCD is to robustify the location and scatter matrix parameters. Since these parameters are highly influenced by the presence of outliers, the DetMCD is a newly highly robust algorithm, where it is constructed to overcome the outlier's problem. DetMCD precedes the non-random subsets, which computes a small number of deterministic initial estimators and followed by concentration steps. Here, we are going to compare the DetMCD algorithm based on two groups of estimators - one with original five Huberts' estimators and the other five new estimators. The determinant values of these estimators are observed to evaluate the performance via several cases.
20 CFR 410.620 - Notice of initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Notice of initial determination. 410.620 Section 410.620 Employees' Benefits SOCIAL SECURITY ADMINISTRATION FEDERAL COAL MINE HEALTH AND SAFETY ACT..., Administrative Review, Finality of Decisions, and Representation of Parties § 410.620 Notice of...
20 CFR 410.621 - Effect of initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Effect of initial determination. 410.621 Section 410.621 Employees' Benefits SOCIAL SECURITY ADMINISTRATION FEDERAL COAL MINE HEALTH AND SAFETY ACT..., Administrative Review, Finality of Decisions, and Representation of Parties § 410.621 Effect of...
42 CFR 405.704 - Actions which are initial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 42 Public Health 2 2010-10-01 2010-10-01 false Actions which are initial determinations. 405.704 Section 405.704 Public Health CENTERS FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH AND HUMAN SERVICES MEDICARE PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED Reconsiderations and...
42 CFR 405.704 - Actions which are initial determinations.
Code of Federal Regulations, 2011 CFR
2011-10-01
... 42 Public Health 2 2011-10-01 2011-10-01 false Actions which are initial determinations. 405.704 Section 405.704 Public Health CENTERS FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH AND HUMAN SERVICES MEDICARE PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED Reconsiderations and...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2014 CFR
2014-10-01
... remittance advice (RA) notice is the notice of initial determination sent to providers and suppliers that accept assignment. The electronic RA must comply with the format and content requirements of the standard... (HIPAA) and related CMS manual instructions. When a paper RA is mailed, it must comply with CMS...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2011 CFR
2011-10-01
... remittance advice (RA) notice is the notice of initial determination sent to providers and suppliers that accept assignment. The electronic RA must comply with the format and content requirements of the standard... (HIPAA) and related CMS manual instructions. When a paper RA is mailed, it must comply with CMS...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2012 CFR
2012-10-01
... remittance advice (RA) notice is the notice of initial determination sent to providers and suppliers that accept assignment. The electronic RA must comply with the format and content requirements of the standard... (HIPAA) and related CMS manual instructions. When a paper RA is mailed, it must comply with CMS...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2013 CFR
2013-10-01
... remittance advice (RA) notice is the notice of initial determination sent to providers and suppliers that accept assignment. The electronic RA must comply with the format and content requirements of the standard... (HIPAA) and related CMS manual instructions. When a paper RA is mailed, it must comply with CMS...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2010 CFR
2010-10-01
... remittance advice (RA) notice is the notice of initial determination sent to providers and suppliers that accept assignment. The electronic RA must comply with the format and content requirements of the standard... (HIPAA) and related CMS manual instructions. When a paper RA is mailed, it must comply with CMS...
Investigating On-Orbit Attitude Determination Anomalies for the Solar Dynamics Observatory Mission
NASA Technical Reports Server (NTRS)
Vess, Melissa F.; Starin, Scott R.; Chia-Kuo, Alice Liu
2011-01-01
The Solar Dynamics Observatory (SDO) was launched on February 11, 2010 from Kennedy Space Center on an Atlas V launch vehicle into a geosynchronous transfer orbit. SDO carries a suite of three scientific instruments, whose observations are intended to promote a more complete understanding of the Sun and its effects on the Earth's environment. After a successful launch, separation, and initial Sun acquisition, the launch and flight operations teams dove into a commissioning campaign that included, among other things, checkout and calibration of the fine attitude sensors and checkout of the Kalman filter (KF) and the spacecraft s inertial pointing and science control modes. In addition, initial calibration of the science instruments was also accomplished. During that process of KF and controller checkout, several interesting observations were noticed and investigated. The SDO fine attitude sensors consist of one Adcole Digital Sun Sensor (DSS), two Galileo Avionica (GA) quaternion-output Star Trackers (STs), and three Kearfott Two-Axis Rate Assemblies (hereafter called inertial reference units, or IRUs). Initial checkout of the fine attitude sensors indicated that all sensors appeared to be functioning properly. Initial calibration maneuvers were planned and executed to update scale factors, drift rate biases, and alignments of the IRUs. After updating the IRU parameters, the KF was initialized and quickly reached convergence. Over the next few hours, it became apparent that there was an oscillation in the sensor residuals and the KF estimation of the IRU bias. A concentrated investigation ensued to determine the cause of the oscillations, their effect on mission requirements, and how to mitigate them. The ensuing analysis determined that the oscillations seen were, in fact, due to an oscillation in the IRU biases. The low frequencies of the oscillations passed through the KF, were well within the controller bandwidth, and therefore the spacecraft was actually following the oscillating biases, resulting in movement of the spacecraft on the order of plus or minus 20 arcsec. Though this level of error met the ACS attitude knowledge requirement of [35, 70, 70] arcsec, 3 sigma, the desire of the ACS and instrument teams was to remove as much of the oscillation as possible. The Kearfott IRUs have an internal temperature controller, designed to maintain the IRU temperature at a constant temperature of approximately 70 C, thus minimizing the change in the bias drift and scale factors of the mechanical gyros. During ground testing of the observatory, it was discovered that the 83-Hz control cycle of the IRU heaters put a tremendous amount of stress on the spacecraft battery. Analysis by the power systems team indicated that the constant charge/discharge on the battery due to the IRU thermal control cycle could potentially limit the life of the battery. After much analysis, the decision was made not to run the internal IRU heaters. Analysis of on orbit data revealed that the oscillations in the IRU bias had a connection to the temperature of the IRU; changes in IRU temperature resulted in changes in the amplitude and period of the IRU biases. Several mitigating solutions were investigated, the result of which was to tune the KF with larger IRU noise assumptions which allows the KF to follow and correct for the time-varying IRU biases.
Landsat-4 (TDRSS-user) orbit determination using batch least-squares and sequential methods
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1992-01-01
TDRSS user orbit determination is analyzed using a batch least-squares method and a sequential estimation method. It was found that in the batch least-squares method analysis, the orbit determination consistency for Landsat-4, which was heavily tracked by TDRSS during January 1991, was about 4 meters in the rms overlap comparisons and about 6 meters in the maximum position differences in overlap comparisons. The consistency was about 10 to 30 meters in the 3 sigma state error covariance function in the sequential method analysis. As a measure of consistency, the first residual of each pass was within the 3 sigma bound in the residual space.
Landsat-4 (TDRSS-user) orbit determination using batch least-squares and sequential methods
NASA Astrophysics Data System (ADS)
Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1992-08-01
TDRSS user orbit determination is analyzed using a batch least-squares method and a sequential estimation method. It was found that in the batch least-squares method analysis, the orbit determination consistency for Landsat-4, which was heavily tracked by TDRSS during January 1991, was about 4 meters in the rms overlap comparisons and about 6 meters in the maximum position differences in overlap comparisons. The consistency was about 10 to 30 meters in the 3 sigma state error covariance function in the sequential method analysis. As a measure of consistency, the first residual of each pass was within the 3 sigma bound in the residual space.
NASA Astrophysics Data System (ADS)
JÃ¤ggi, Adrian; Bock, Heike; Dach, Rolf; Montenbruck, Oliver; Hugentobler, Urs; Beutler, Gerhard
Absolute phase patterns for GNSS receiver and transmitter antennas are adopted in the processing standards of the International GNSS Service (IGS) since November 5, 2006 (GPS week 1400). The new antenna modeling is based on robot-calibrations for a number of terrestrial receiver antennas. Compatible antenna models are derived for the remaining terrestrial receiver antennas and the GNSS satellite antennas. However, consistent receiver antenna patterns are not available for many space missions equipped with onboard GPS sensor systems. Recently, nominal phase patterns obtained with a robotic measurement system in a field campaign have been made available for the antenna/chokering combination deployed on the CHAMP, GRACE, and TerraSAR-X satellites. We use the final product line from the Center for Orbit Determination in Europe (CODE: analysis center of the IGS) together with GPS data of the aforementioned low Earth orbiters (LEOs) for the year 2007 to assess the impact of nominal phase patterns on reduced-dynamic and kinematic LEO orbits computed by different and independent software packages (Bernese GPS software, GPS High-precision Orbit Determination Software Tools). In the actual spacecraft environment, however, pronounced phase center distortions may be encountered in addition due to multipath or cross-talk effects, which makes an additional in-flight calibration of LEO receiver antennas desirable. We compare methods for the in-flight derivation of empirical phase pattern corrections and discuss their relevance and applicability for precise orbit determination.
Accurate orbit determination strategies for the tracking and data relay satellites
NASA Technical Reports Server (NTRS)
Oza, D. H.; Bolvin, D. T.; Lorah, J. M.; Lee, T.; Doll, C. E.
1995-01-01
The National Aeronautics and Space Administration (NASA) has developed the Tracking and Data Relay Satellite (TDRS) System (TDRSS) for tracking and communications support of low Earth-orbiting satellites. TDRSS has the operational capability of providing 85% coverage for TDRSS-user spacecraft. TDRSS currently consists of five geosynchronous spacecraft and the White Sands Complex (WSC) at White Sands, New Mexico. The Bilateration Ranging Transponder System (BRTS) provides range and Doppler measurements for each TDRS. The ground-based BRTS transponders are tracked as if they were TDRSS-user spacecraft. Since the positions of the BRTS transponders are known, their radiometric tracking measurements can be used to provide a well-determined ephemeris for the TDRS spacecraft. For high-accuracy orbit determination of a TDRSS user, such as the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft, high-accuracy TDRS orbits are required. This paper reports on successive refinements in improved techniques and procedures leading to more accurate TDRS orbit determination strategies using the Goddard Trajectory Determination System (GTDS). These strategies range from the standard operational solution using only the BRTS tracking measurements to a sophisticated iterative process involving several successive simultaneous solutions for multiple TDRSs and a TDRSS-user spacecraft. Results are presented for GTDS-generated TDRS ephemerides produced in simultaneous solutions with the TOPEX/Poseidon spacecraft. Strategies with different user spacecraft, as well as schemes for recovering accurate TDRS orbits following a TDRS maneuver, are also presented. In addition, a comprehensive assessment and evaluation of alternative strategies for TDRS orbit determination, excluding BRTS tracking measurements, are presented.
Precise orbit determination of BeiDou constellation based on BETS and MGEX network
Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu
2014-01-01
Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20â€…cm and 14â€…cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10â€…cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved. PMID:24733025
Precise orbit determination of BeiDou constellation based on BETS and MGEX network.
Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu
2014-01-01
Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20â€…cm and 14â€…cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10â€…cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved. PMID:24733025
Precise orbit determination of BeiDou constellation based on BETS and MGEX network
NASA Astrophysics Data System (ADS)
Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu
2014-04-01
Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20 cm and 14 cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10 cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved.
Orbit Determination Using SLR Data for STSAT-2C: Short-arc Analysis
NASA Astrophysics Data System (ADS)
Kim, Young-Rok; Park, Eunseo; Kucharski, Daniel; Lim, Hyung-Chul
2015-09-01
In this study, we present the results of orbit determination (OD) using satellite laser ranging (SLR) data for the Science and Technology Satellite (STSAT)-2C by a short-arc analysis. For SLR data processing, the NASA/GSFC GEODYN II software with one year (2013/04 - 2014/04) of normal point observations is used. As there is only an extremely small quantity of SLR observations of STSAT-2C and they are sparsely distribution, the selection of the arc length and the estimation intervals for the atmospheric drag coefficients and the empirical acceleration parameters was made on an arc-to-arc basis. For orbit quality assessment, the post-fit residuals of each short-arc and orbit overlaps of arcs are investigated. The OD results show that the weighted root mean square post-fit residuals of short-arcs are less than 1 cm, and the average 1-day orbit overlaps are superior to 50/600/900 m for the radial/cross-track/along-track components. These results demonstrate that OD for STSAT-2C was successfully achieved with cm-level range precision. However its orbit quality did not reach the same level due to the availability of few and sparse measurement conditions. From a mission analysis viewpoint, obtaining the results of OD for STSAT-2C is significant for generating enhanced orbit predictions for more frequent tracking.
NASA Technical Reports Server (NTRS)
Wu, Jiun-Tsong; Wu, Sien-Chong
1992-01-01
A method to determine satellite orbits using tracking data and a priori gravitational field is described. The a priori constraint on the orbit dynamics is determined by the covariance matrix of the spherical harmonic coefficients for the gravity model, so that the optimal combination of the measurements and gravitational field is achieved. A set of bin parameters is introduced to represent the perturbation of the gravitational field on the position of the satellite orbit. The covariance matrix of a conventional gravity model is transformed into that for the bin parameters by the variational partial derivatives. The covariance matrices of the bin parameters and the epoch state are combined to form the covariance matrix of the satellite positions at the measurement times. The combined matrix is used as the a priori information to estimate the satellite positions with measurements.
Four Methods for Determining Intermediate Perturbed Orbits from Three Observations: A Comparison
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2015-03-01
Theoretical and numerical comparison of four methods for determining the orbit of a small celestial body from three measurements of its angular coordinates at three time moments is provided. The methods are intended for constructing intermediate orbits considering most of perturbations in motion of the examined body. Two methods are based on the solutions of the differential equations of motion and on their second derivatives in the form of series in terms of powers of time intervals (the Herrick-Gibbs approach), and the two others are based on the solutions for some intermediate perturbed motions in the closed form, without their representation in the form of series (the approach of the author). A dependence of methodic errors on the length of the reference time interval determined by the moments of observation beginning and ending is investigated. By way of examples, results of calculation of the orbit of the Apophis asteroid are presented.
NASA Technical Reports Server (NTRS)
Mardirossian, H.; Heuerman, K.; Beri, A.; Samii, M. V.; Doll, C. E.
1989-01-01
The Flight Dynamics Facility (FDF) at Goddard Space Flight Center (GSFC) provides spacecraft trajectory determination for a wide variety of National Aeronautics and Space Administration (NASA)-supported satellite missions, using the Tracking Data Relay Satellite System (TDRSS) and Ground Spaceflight and Tracking Data Network (GSTDN). To take advantage of computerized decision making processes that can be used in spacecraft navigation, the Orbit Determination Automation System (ODAS) was designed, developed, and implemented as a prototype system to automate orbit determination (OD) and orbit quality assurance (QA) functions performed by orbit operations. Based on a machine-resident generic schedule and predetermined mission-dependent QA criteria, ODAS autonomously activates an interface with the existing trajectory determination system using a batch least-squares differential correction algorithm to perform the basic OD functions. The computational parameters determined during the OD are processed to make computerized decisions regarding QA, and a controlled recovery process isactivated when the criteria are not satisfied. The complete cycle is autonomous and continuous. ODAS was extensively tested for performance under conditions resembling actual operational conditions and found to be effective and reliable for extended autonomous OD. Details of the system structure and function are discussed, and test results are presented.
NASA Technical Reports Server (NTRS)
Mardirossian, H.; Beri, A. C.; Doll, C. E.
1990-01-01
The Flight Dynamics Facility (FDF) at Goddard Space Flight Center (GSFC) provides spacecraft trajectory determination for a wide variety of National Aeronautics and Space Administration (NASA)-supported satellite missions, using the Tracking Data Relay Satellite System (TDRSS) and Ground Spaceflight and Tracking Data Network (GSTDN). To take advantage of computerized decision making processes that can be used in spacecraft navigation, the Orbit Determination Automation System (ODAS) was designed, developed, and implemented as a prototype system to automate orbit determination (OD) and orbit quality assurance (QA) functions performed by orbit operations. Based on a machine-resident generic schedule and predetermined mission-dependent QA criteria, ODAS autonomously activates an interface with the existing trajectory determination system using a batch least-squares differential correction algorithm to perform the basic OD functions. The computational parameters determined during the OD are processed to make computerized decisions regarding QA, and a controlled recovery process is activated when the criteria are not satisfied. The complete cycle is autonomous and continuous. ODAS was extensively tested for performance under conditions resembling actual operational conditions and found to be effective and reliable for extended autonomous OD. Details of the system structure and function are discussed, and test results are presented.
Mitigation of ionospheric scintillation effects in kinematic LEO precise orbit determination
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-GÃ¼rr, Torsten
2015-04-01
Kinematic orbit determination for Low Earth Orbiting satellites is one of the core elements in gravity field recovery from GNSS tracked satellites. The accuracy of the kinematic orbit positions directly determines the achievable accuracy in terms of gravity field results. We apply a precise point positioning approach based on raw GNSS observations, without using any linear combinations. This method requires to take every effect directly into account, as non of the effects is eliminated by forming differences or linear combinations. For example, the ionospheric influence is taken into account by estimating the slant TEC, including higher order terms and corrections for ionospheric bending. Our approach preserves the original high measurement accuracy of the phase observations. The remaining factors reducing the achieved accuracy are not or incorrectly modeled systematic effects. The GOCE mission revealed one of these systematic effects: ionospheric scintillations. These are small and short term irregularities in the Earth's ionosphere which cause errors in GNSS observations. GOCE gravity field results showed a huge systematic effect along the geomagnetic equator. GOCE was flying in a sun-synchronous dusk-dawn orbit, which means that the satellite orbit is nearly stationary with respect to the Earth's ionosphere. As it is hardly possible to realistically model ionospheric irregularities they can not be corrected from the raw observations. We introduce an observation weighting method based on the rate of TEC index to reduce the influence of observations affected by ionospheric scintillations. This weighting scheme in combination with variance component estimation greatly reduces the influence of ionospheric scintillation on the kinematic orbit and in turn also on the gravity field result. We will show that by using the introduced weighting scheme the error in GOCE kinematic orbits is almost removed, without removing observations.
Precise Orbit Determination Of Low Earth Satellites At AIUB Using GPS And SLR Data
NASA Astrophysics Data System (ADS)
Jaggi, A.; Bock, H.; Thaller, D.; Sosnica, K.; Meyer, U.; Baumann, C.; Dach, R.
2013-12-01
An ever increasing number of low Earth orbiting (LEO) satellites is, or will be, equipped with retro-reflectors for Satellite Laser Ranging (SLR) and on-board receivers to collect observations from Global Navigation Satellite Systems (GNSS) such as the Global Positioning System (GPS) and the Russian GLONASS and the European Galileo systems in the future. At the Astronomical Institute of the University of Bern (AIUB) LEO precise orbit determination (POD) using either GPS or SLR data is performed for a wide range of applications for satellites at different altitudes. For this purpose the classical numerical integration techniques, as also used for dynamic orbit determination of satellites at high altitudes, are extended by pseudo-stochastic orbit modeling techniques to efficiently cope with potential force model deficiencies for satellites at low altitudes. Accuracies of better than 2 cm may be achieved by pseudo-stochastic orbit modeling for satellites at very low altitudes such as for the GPS-based POD of the Gravity field and steady-state Ocean Circulation Explorer (GOCE).
NASA Astrophysics Data System (ADS)
Flohrer, Tim; Beutler, Gerhard; Schildknecht, Thomas
Surveys for space debris aim at improving our knowledge of the space debris population. The survey results are fed either into space debris environment models or into orbital element catalogue of space debris objects, depending on whether the objects should be tracked later on. In both cases first orbit determination is a crucial step. The determined orbital elements together with estimated of the magnitude of the object allow first estimates of the object size. The orbital elements are also the central input to validate and improve the space debris environment models. Considering catalogue build-up and maintenance, the results from the first orbit determination set important constraints for the correlation of the new object with the catalogue, the re-acquisition and finally the identification of the object. In the case of space debris surveys first orbits must be determined from very short arcs of observations, which are due to the limits of the current sensor technology, in particular due to the limited field-of-view diameter of the telescopes used for optical surveys. The quality of the first orbits is in any case a function of the measurement accuracy. We present an approach for first orbit determination, which is derived from the boundary value method implemented in the CelMech program system (Beutler, 2005). The approach considers angular measurements, as derived from optical observation systems. The CelMech module ORBDET was generalized to perform a two-dimensional search by systematically varying the topocentric ranges at the boundary epochs of the observed arc. The search aims at identifying local minima of a least-square adjustment of all available observations using a truncated Taylor series to approximate the particular solution of the equation of motion for the debris considered. We apply this method to observations from ongoing space debris surveys of ESA using the 1-m telescope at the Optical Ground Station in Tenerife. Furthermore we apply this approach to a proposed space-based optical observation system. Last, but not least we consider the inclusion of range observations in our algorithm. Beutler G (2005) Methods of Celestial Mechanics. Springer, Berlin, Heidelberg, New York
TOPEX orbit determination using GPS signals plus a sidetone ranging system
NASA Technical Reports Server (NTRS)
Bender, P. L.; Larden, D. R.
1982-01-01
The GPS orbit determination was studied to see how well the radial coordinate for altimeter satellites such as TOPEX could be found by on board measurements of GPS signals, including the reconstructed carrier phase. The inclusion on altimeter satellites of an additional high accuracy tracking system is recommended. It is suggested that a sidetone ranging system is used in conjunction with TRANET 2 beacons.
Researches on the Orbit Determination and Positioning of the Chinese Lunar Exploration Program
NASA Astrophysics Data System (ADS)
Li, P. J.
2015-07-01
This dissertation studies the precise orbit determination (POD) and positioning of the Chinese lunar exploration spacecraft, emphasizing the variety of VLBI (very long baseline interferometry) technologies applied for the deep-space exploration, and their contributions to the methods and accuracies of the precise orbit determination and positioning. In summary, the main contents are as following: In this work, using the real-time data measured by the CE-2 (Chang'E-2) detector, the accuracy of orbit determination is analyzed for the domestic lunar probe under the present condition, and the role played by the VLBI tracking data is particularly reassessed through the precision orbit determination experiments for CE-2. The experiments of the short-arc orbit determination for the lunar probe show that the combination of the ranging and VLBI data with the arc of 15 minutes is able to improve the accuracy by 1-1.5 order of magnitude, compared to the cases for only using the ranging data with the arc of 3 hours. The orbital accuracy is assessed through the orbital overlapping analysis, and the results show that the VLBI data is able to contribute to the CE-2's long-arc POD especially in the along-track and orbital normal directions. For the CE-2's 100 kmÃ— 100 km lunar orbit, the position errors are better than 30 meters, and for the CE-2's 15 kmÃ— 100 km orbit, the position errors are better than 45 meters. The observational data with the delta differential one-way ranging (Î” DOR) from the CE-2's X-band monitoring and control system experimental are analyzed. It is concluded that the accuracy of Î” DOR delay is dramatically improved with the noise level better than 0.1 ns, and the systematic errors are well calibrated. Although it is unable to support the development of an independent lunar gravity model, the tracking data of CE-2 provided the evaluations of different lunar gravity models through POD, and the accuracies are examined in terms of orbit-to-orbit solution differences for several gravity models. It is found that for the 100 kmÃ— 100 km lunar orbit, with a degree and order expansion up to 165, the JPL's gravity model LP165P does not show noticeable improvement over Japan's SGM series models (100Ã— 100), but for the 15 kmÃ— 100 km lunar orbit, a higher degree-order model can significantly improve the orbit accuracy. After accomplished its nominal mission, CE-2 launched its extended missions, which involving the L2 mission and the 4179 Toutatis mission. During the flight of the extended missions, the regime offers very little dynamics thus requires an extensive amount of time and tracking data in order to attain a solution. The overlap errors are computed, and it is indicated that the use of VLBI measurements is able to increase the accuracy and reduce the total amount of tracking time. An orbit determination method based on the polynomial fitting is proposed for the CE-3's planned lunar soft landing mission. In this method, spacecraft's dynamic modeling is not necessary, and its noise reduction is expected to be better than that of the point positioning method by making full use of all-arc observational data. The simulation experiments and real data processing showed that the optimal description of the CE-1's free-fall landing trajectory is a set of five-order polynomial functions for each of the position components as well as velocity components in J2000.0. The combination of the VLBI delay, the delay rate data, and the USB (united S-band) ranging data significantly improved the accuracy than the use of USB data alone. In order to determine the position for the CE-3's Lunar Lander, a kinematic statistical method is proposed. This method uses both ranging and VLBI measurements to the lander for a continuous arc, combing with precise knowledge about the motion of the moon as provided by planetary ephemeris, to estimate the lander's position on the lunar surface with high accuracy. Application of the lunar digital elevation model (DEM) as constraints in the lander positioning is helpful. The positioning method for the traverse of lunar rover is also investigated. The integration of delay-rate method is able to achieve higher precise positioning results than the point positioning method. This method provides a wide application of the VLBI data. In the automated sample return mission, the lunar orbit rendezvous and docking are involved. Precise orbit determination using the same-beam VLBI (SBI) measurement for two spacecraft at the same time is analyzed. The simulation results showed that the SBI data is able to improve the absolute and relative orbit accuracy for two targets by 1-2 orders of magnitude. In order to verify the simulation results and test the two-target POD software developed by SHAO (Shanghai Astronomical Observatory), the real SBI data of the SELENE (Selenological and Engineering Explorer) are processed. The POD results for the Rstar and the Vstar showed that the combination of SBI data could significantly improve the accuracy for the two spacecraft, especially for the Vstar with less ranging data, and the POD accuracy is improved by approximate one order of magnitude to the POD accuracy of the Rstar.
NASA Technical Reports Server (NTRS)
Keckler, C. R.; Kibler, K. S.; Powell, L. F.
1979-01-01
A high fidelity simulation of the annular suspension and pointing system (ASPS), its payload, and the shuttle orbiter was used to define the worst case orientations of the ASPS and its payload for the various vehicle disturbances, and to determine the performance capability of the ASPS under these conditions. The most demanding and largest proposed payload, the Solar Optical Telescope was selected for study. It was found that, in all cases, the ASPS more than satisfied the payload's requirements. It is concluded that, to satisfy facility class payload requirements, the ASPS or a shuttle orbiter free-drift mode (control system off) should be utilized.
NASA Technical Reports Server (NTRS)
Vigue, Y.; Lichten, S. M.; Muellerschoen, R. J.; Blewitt, G.; Heflin, M. B.
1993-01-01
Data collected from a worldwide 1992 experiment were processed at JPL to determine precise orbits for the satellites of the Global Positioning System (GPS). A filtering technique was tested to improve modeling of solar-radiation pressure force parameters for GPS satellites. The new approach improves orbit quality for eclipsing satellites by a factor of two, with typical results in the 25- to 50-cm range. The resultant GPS-based estimates for geocentric coordinates of the tracking sites, which include the three DSN sites, are accurate to 2 to 8 cm, roughly equivalent to 3 to 10 nrad of angular measure.
The effects of geopotential resonance on orbit determination for Landsat-4
NASA Technical Reports Server (NTRS)
Hoge, S. L.; Casteel, D. O.; Phenneger, M. C.; Smith, E. A.
1988-01-01
Analysis is presented demonstrating improved performance for Landsat-4 orbit determination using the Goddard Trajectory Determination System with an adjusted Goddard Earth Model-9 (GEM-9) for geopotential coefficients of the 15th degree and order. The orbital state is estimated along with the sine and cosine coefficients of degree and order 15, (C, S) sub 15,15. The estimates are made for two 5-day intervals of range and doppler data, primarily from the Tracking and Data Relay Satellite, during a period of low solar activity in January 1987. The average values of the estimated coefficients (C, S) sub 15,15 are used to modify the GEM-9 model, and orbit determination performance is tested on 17 consecutive 34-hour operational tracking data arcs in January 1987. Significant reductions in the mean values and standard deviations of the along-track position difference and the drag model scaling parameter from solution to solution are observed. The approach is guided by the shallow resonance theory of geopotential orbit perturbations.
A multi-satellite orbit determination problem in a parallel processing environment
NASA Technical Reports Server (NTRS)
Deakyne, M. S.; Anderle, R. J.
1988-01-01
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.
A new procedure for orbit determination based on three lines of sight: Angles only
NASA Astrophysics Data System (ADS)
Gooding, R. H.
1993-04-01
A new procedure has been developed for the general solution of the minimal angles-only problem in which an orbit is determined from three line-of-sight observations. The basis of the approach is a higher-order Newton correction of the assumed values for two of the unknown ranges, appeal being made to the author's (published) universal solution of Lambert's orbital boundary-value problem. The new procedure is free of the inherent limitations of the traditional methods of Laplace and Gauss, these methods being outlined in a summary of previous approaches to this classical problem. In particular, the observations are permitted to span several revolutions when the orbit is elliptic; multiple solutions can be obtained; and there is no restriction on the configuration of the three observing sites. The procedure has been carefully tested, some of the examples being taken from the literature. A number of test problems have been solved that would have failed by existing methods.
The determination of maximum deep space station slew rates for a high Earth orbiter
NASA Technical Reports Server (NTRS)
Estefan, J. A.
1990-01-01
As developing national and international space ventures, which seek to employ NASA's Deep Space Network (DSN) for tracking and data acquisition, evolve, it is essential for navigation and tracking system analysts to evaluate the operational capability of Deep Space Station antennas. To commission the DSN for use in tracking a highly eccentric Earth orbiter could quite possibly yield the greatest challenges in terms of slewing capability; certainly more so than with a deep-space probe. The focus here is on the determination of the maximum slew rates needed to track a specific high Earth orbiter, namely the Japanese MUSES-B spacecraft of the Very Long Baseline Interferometry Space Observatory Program. The results suggest that DSN 34-m antennas are capable of meeting the slew rate requirements for the nominal MUSES-B orbital geometries currently being considered.
TOPEX/POSEIDON operational orbit determination results using global positioning satellites
NASA Technical Reports Server (NTRS)
Guinn, J.; Jee, J.; Wolff, P.; Lagattuta, F.; Drain, T.; Sierra, V.
1994-01-01
Results of operational orbit determination, performed as part of the TOPEX/POSEIDON (T/P) Global Positioning System (GPS) demonstration experiment, are presented in this article. Elements of this experiment include the GPS satellite constellation, the GPS demonstration receiver on board T/P, six ground GPS receivers, the GPS Data Handling Facility, and the GPS Data Processing Facility (GDPF). Carrier phase and P-code pseudorange measurements from up to 24 GPS satellites to the seven GPS receivers are processed simultaneously with the GDPF software MIRAGE to produce orbit solutions of T/P and the GPS satellites. Daily solutions yield subdecimeter radial accuracies compared to other GPS, LASER, and DORIS precision orbit solutions.
Determining Mars parking orbits which ensure in-plane arrival and departure burns
NASA Technical Reports Server (NTRS)
Desai, Prasun N.; Buglia, James J.
1992-01-01
A numerical method to find suitable Mars parking orbits is developed which takes into account geometries associated with the asymptotes, along with the nodal precession caused by the oblateness of Mars. A selected orbital plane which contains the arrival asymptote precesses through the stay time to the plane also containing the departure asymptote. The parking orbit is co-planar with both the arrival and departure asymptotes and only in-plane burns are required at both Mars arrival and departure. The need for a plane change at Mars departure to achieve the proper velocity vector for earth return is eliminated. The method requires very little computation time to determine a set of all possible inclinations and right ascensions of the ascending nodes.
NASA Astrophysics Data System (ADS)
Couhert, Alexandre
The reference Ocean Surface Topography Mission/Jason-2 satellite (CNES/NASA) has been in orbit for six years (since June 2008). It extends the continuous record of highly accurate sea surface height measurements begun in 1992 by the Topex/Poseidon mission and continued in 2001 by the Jason-1 mission. The complementary missions CryoSat-2 (ESA), HY-2A (CNSA) and SARAL/AltiKa (CNES/ISRO), with lower altitudes and higher inclinations, were launched in April 2010, August 2011 and February 2013, respectively. Although the three last satellites fly in different orbits, they contribute to the altimeter constellation while enhancing the global coverage. The CNES Precision Orbit Determination (POD) Group delivers precise and homogeneous orbit solutions for these independent altimeter missions. The focus of this talk will be on the long-term stability of the orbit time series for mean sea level applications on a regional scale. We discuss various issues related to the assessment of radial orbit error trends; in particular orbit errors dependant on the tracking technique, the reference frame accuracy and stability, the modeling of the temporal variations of the geopotential. Strategies are then explored to meet a 1 mm/y radial orbit stability over decadal periods at regional scales, and the challenge of evaluating such an improvement is discussed.
NASA Astrophysics Data System (ADS)
Keating, G. M.; Bougher, S. W.; Theriot, M. E.; Tolson, R. H.; Blanchard, R. C.; Zurek, R. W.; Forbes, J. M.; Murphy, J.
2006-12-01
Designed for aerobraking, Mars Reconnaissance Orbiter (MRO) launched on August 12, 2005, achieved Mars Orbital Insertion (MOI), March 10, 2006, and successfully completed aerobraking on August 30, 2006. Atmospheric density decreases exponentially with increasing height. By small propulsive adjustments of the apoapsis orbital velocity, periapsis altitude was fine tuned to the density surface that safely used the atmosphere of Mars to aerobrake over 445 orbits, providing 890 vertical structures. MRO periapsis precesses from near the South Pole at 6pm LST to near the equator at 3am LST. Meanwhile, apoapsis is brought dramatically from 40,000km at MOI to 480 km at aerobraking completion (ABX). Without aerobraking this would have required an additional 400kg of fuel. After ABX, two small propulsive orbital adjustment maneuvers September 5, 2006 and September 11, 2006 established the final Primary Science Orbit (PSO). Each of the 445 aerobraking orbits provides, a pair of vertical structures inbound toward periapsis and outbound from periapsis, with a distribution of density, scale heights, temperatures, and pressures along the orbital path, providing key in situ insight into various upper atmosphere (> 100 km) processes. One of the major questions for scientists studying Mars is: Where did the water go? Honeywell's substantially improved electronics package for its IMU (QA-2000 accelerometer, gyro, electronics) maximized accelerometer sensitivities at the requests of The George Washington University, JPL, and Lockheed Martin. The improved accelerometer sensitivities allowed density measurements to exceed 200km, at least 40 km higher than with Mars Odyssey (MO). This extends vertical structures from MRO into the neutral lower exosphere, a region where various processes may allow atmospheric gasses to escape. Over the eons, water may have been lost in both the lower atmosphere and the upper atmosphere, thus the water balance throughout the entire atmosphere from subsurface to exosphere may be equally critical. Comparisons of accelerometer data from Mars Global Surveyor (MGS), MO and MRO will help characterize key temporal and spatial cycles. During the Odyssey Aerobraking we discovered a very strong winter polar warming near 100km, where temperatures were found to be up to 100K higher than expected near the North Pole. However, with MRO we detected only a very weak winter polar warming at the South Pole. It is expected that the polar warming results from cross equatorial meridional flow from the summer hemisphere into the winter hemisphere with adiabatic heating near the winter pole. The discovery from MRO of a very weak winter warming near aphelion in the southern winter polar region compared to the very strong winter warming near perihelion in the northern winter polar region is apparently due to a weaker input of solar energy into the meridional circulation resulting in less adiabatic heating near aphelion in the winter polar region. Results are also shown of global scale measurements of non- migrating tides and of global density and temperature distributions.
An initial comparative assessment of orbital and terrestrial central power systems
NASA Technical Reports Server (NTRS)
Caputo, R.
1977-01-01
A silicon photovoltaic orbital power system, which is constructed from an earth source of materials, is compared to likely terrestrial (fossil, nuclear, and solar) approaches to central power generation around the year 2000. A total social framework is used that considers not only the projection of commercial economics (direct or in internal costs), but also considers external impacts such as research and development investment, health impacts, resource requirements, environment effects, and other social costs.
The GLAS Algorithm Theoretical Basis Document for Precision Orbit Determination (POD)
NASA Technical Reports Server (NTRS)
Rim, Hyung Jin; Yoon, S. P.; Schultz, Bob E.
2013-01-01
The Geoscience Laser Altimeter System (GLAS) was the sole instrument for NASA's Ice, Cloud and land Elevation Satellite (ICESat) laser altimetry mission. The primary purpose of the ICESat mission was to make ice sheet elevation measurements of the polar regions. Additional goals were to measure the global distribution of clouds and aerosols and to map sea ice, land topography and vegetation. ICESat was the benchmark Earth Observing System (EOS) mission to be used to determine the mass balance of the ice sheets, as well as for providing cloud property information, especially for stratospheric clouds common over polar areas. The GLAS instrument operated from 2003 to 2009 and provided multi-year elevation data needed to determine changes in sea ice freeboard, land topography and vegetation around the globe, in addition to elevation changes of the Greenland and Antarctic ice sheets. This document describes the Precision Orbit Determination (POD) algorithm for the ICESat mission. The problem of determining an accurate ephemeris for an orbiting satellite involves estimating the position and velocity of the satellite from a sequence of observations. The ICESatGLAS elevation measurements must be very accurately geolocated, combining precise orbit information with precision pointing information. The ICESat mission POD requirement states that the position of the instrument should be determined with an accuracy of 5 and 20 cm (1-s) in radial and horizontal components, respectively, to meet the science requirements for determining elevation change.
NASA Technical Reports Server (NTRS)
Doll, C. E.; Gramling, C. J.; Oza, D. H.; Radomski, M. S.
1990-01-01
The results of a study to analyze the dependence of TDRSS user spacecraft orbit determination consistencies on varying tracking schedules are presented. In this study, the TDRS-East orbit determination results obtained utilizing Bilateration Ranging Transponder System data were evaluated. Six state parameters, three position and three velocity components and the solar radiation pressure coefficient, are estimated for TDRS-East. It is concluded that, in order to achieve high-precision orbit determination, the tracking coverage should not fall below 10 minutes every two orbits as decreasing it to every four orbits will significantly degrade the accuracy; present state-of-the-art consistency in orbit determination using TDRSS tracking is approximately 15 to 20 meters.
The Role of GRAIL Orbit Determination in Preprocessing of Gravity Science Measurements
NASA Technical Reports Server (NTRS)
Kruizinga, Gerhard; Asmar, Sami; Fahnestock, Eugene; Harvey, Nate; Kahan, Daniel; Konopliv, Alex; Oudrhiri, Kamal; Paik, Meegyeong; Park, Ryan; Strekalov, Dmitry; Watkins, Michael; Yuan, Dah-Ning
2013-01-01
The Gravity Recovery And Interior Laboratory (GRAIL) mission has constructed a lunar gravity field with unprecedented uniform accuracy on the farside and nearside of the Moon. GRAIL lunar gravity field determination begins with preprocessing of the gravity science measurements by applying corrections for time tag error, general relativity, measurement noise and biases. Gravity field determination requires the generation of spacecraft ephemerides of an accuracy not attainable with the pre-GRAIL lunar gravity fields. Therefore, a bootstrapping strategy was developed, iterating between science data preprocessing and lunar gravity field estimation in order to construct sufficiently accurate orbit ephemerides.This paper describes the GRAIL measurements, their dependence on the spacecraft ephemerides and the role of orbit determination in the bootstrapping strategy. Simulation results will be presented that validate the bootstrapping strategy followed by bootstrapping results for flight data, which have led to the latest GRAIL lunar gravity fields.
GPS-Based Navigation and Orbit Determination for the AMSAT Phase 3D Satellite
NASA Technical Reports Server (NTRS)
Davis, George; Carpenter, Russell; Moreau, Michael; Bauer, Frank H.; Long, Anne; Kelbel, David; Martin, Thomas
2002-01-01
This paper summarizes the results of processing GPS data from the AMSAT Phase 3D (AP3) satellite for real-time navigation and post-processed orbit determination experiments. AP3 was launched into a geostationary transfer orbit (GTO) on November 16, 2000 from Kourou, French Guiana, and then was maneuvered into its HEO over the next several months. It carries two Trimble TANS Vector GPS receivers for signal reception at apogee and at perigee. Its spin stabilization mode currently makes it favorable to track GPS satellites from the backside of the constellation while at perigee, and to track GPS satellites from below while at perigee. To date, the experiment has demonstrated that it is feasible to use GPS for navigation and orbit determination in HEO, which will be of great benefit to planned and proposed missions that will utilize such orbits for science observations. It has also shown that there are many important operational considerations to take into account. For example, GPS signals can be tracked above the constellation at altitudes as high as 58000 km, but sufficient amplification of those weak signals is needed. Moreover, GPS receivers can track up to 4 GPS satellites at perigee while moving as fast as 9.8 km/sec, but unless the receiver can maintain lock on the signals long enough, point solutions will be difficult to generate. The spin stabilization of AP3, for example, appears to cause signal levels to fluctuate as other antennas on the satellite block the signals. As a result, its TANS Vectors have been unable to lock on to the GPS signals long enough to down load the broadcast ephemeris and then generate position and velocity solutions. AP3 is currently in its eclipse season, and thus most of the spacecraft subsystems have been powered off. In Spring 2002, they will again be powered up and AP3 will be placed into a three-axis stabilization mode. This will significantly enhance the likelihood that point solutions can be generated, and perhaps more important, that the receiver clock can be synchronized to GPS time. This is extremely important for real-time and post-processed orbit determination, where removal of receiver clock bias from the data time tags is needed, for time-tagging of science observations. Current analysis suggests that the inability to generate point solutions has allowed the TANS Vector clock bias to drift freely, being perhaps as large as 5-7 seconds by October, 2001, thus causing up to 50 km of along-track orbit error. The data collected in May, 2002 while in three-axis stabilized mode should provide a significant improvement in the orbit determination results.
NASA Technical Reports Server (NTRS)
Lindqwister, Ulf J.; Lichten, Stephen M.; Davis, Edgar S.; Theiss, Harold L.
1993-01-01
Topex/Poseidon, a cooperative satellite mission between United States and France, aims to determine global ocean circulation patterns and to study their influence on world climate through precise measurements of sea surface height above the geoid with an on-board altimeter. To achieve the mission science aims, a goal of 13-cm orbit altitude accuracy was set. Topex/Poseidon includes a Global Positioning System (GPS) precise orbit determination (POD) system that has now demonstrated altitude accuracy better than 5 cm. The GPS POD system includes an on-board GPS receiver and a 6-station GPS global tracking network. This paper reviews early GPS results and discusses multi-mission capabilities available from a future enhanced global GPS network, which would provide ground-based geodetic and atmospheric calibrations needed for NASA deep space missions while also supplying tracking data for future low Earth orbiters. Benefits of the enhanced global GPS network include lower operations costs for deep space tracking and many scientific and societal benefits from the low Earth orbiter missions, including improved understanding of ocean circulation, ocean-weather interactions, the El Nino effect, the Earth thermal balance, and weather forecasting.
NASA Astrophysics Data System (ADS)
Soudarin, L.; Capdeville, H.; Lemoine, J.-M.; Schaeffer, P.
2012-04-01
At the end of 2011, the CNES/CLS Analysis Center has entirely re-processed the whole DORIS data set for orbit determination and tracking station coordinate estimation. In addition to SPOT-2, -3, -4, -5, Topex/Poseidon and Envisat, the DORIS/DGXX measurements of Jason-2 and Cryosat-2 are included in the products delivered to the IDS (combined multi-satellite weekly SINEX, orbits in sp3 format). The new processing was motivated by upgrades brought to the GINS/DYNAMO software and the availability of new models. Changes with respect to the previous processing set up for the IDS-3 realization (IDS solution contributing to ITRF2008 computation) are: - a priori reference system defined by DPOD2008 (also used for discontinuities and data rejection) and IERS EOP series aligned on ITRF2008; - trospospheric delays derived from GMF/GPT model; - EIGEN-6S gravity model. Attitude laws implemented in GINS have been revised. A new macro-model tuned by GRGS is now used for Jason-2. The objective of this presentation is to show the impact of this reprocessing on the orbit determination and the terrestrial reference frame. Post-fit residuals, orbit comparison, estimated dynamical parameters are discussed, as well as station positioning performances. Residual signals at draconitic and beta-prime periods are also examined, especially in the geocenter time series.
Period and Orbital Separation determination of a Subdwarf B Pulsator, EC 20117-4014
NASA Astrophysics Data System (ADS)
Otani, Tomomi; Oswalt, Terry
2016-01-01
EC 20117-4014 (V4640 Sgr) is believed to be a binary system consisting of a pulsating subdwarf B star and a F5V star, however the binary period and orbital distance has not been firmly determined. So far, the most promising theory for the origin of subdwarf B (sdB) stars is that they result from binary mass transfer near the Helium Flush stage. We attempted to constrain this evolutional theory by searching for companions and determining periods and orbital separations around sdB pulsators using the Observed-minus-Calculated (O-C) method. A star's position in space will wobble due to the gravitational forces of any companion. If the star is emitting a periodic signal, its orbital motion around the system's center of mass causes periodic changes in the light pulse arrival times. EC 20117-4014 was monitored from 2010-1 using the 0.6m SARA-CT telescope in Cerro Tololo Inter-American Observatory, Chile. After obtaining the O-C diagrams for the star, useful limits on suspected companions' minimum masses and semimajor axes were calculated. In addition, a modeling experiment was performed to investigate the ranges and combinations of possible companion masses and orbits that are consistent with the observational data. Also, the expected radial velocity semi-amplitude for each O-C companion signal was estimated.
An initial comparative assessment of orbital and terrestrial central power systems
NASA Technical Reports Server (NTRS)
Caputo, R.
1977-01-01
Orbital solar power plants, which beam power to earth by microwave, are compared with ground-based solar and conventional baseload power plants. Candidate systems were identified for three types of plants and the selected plant designs were then compared on the basis of economic and social costs. The representative types of plant selected for the comparison are: light water nuclear reactor; turbines using low BTU gas from coal; central receiver with steam turbo-electric conversion and thermal storage; silicon photovoltaic power plant without tracking and including solar concentration and redox battery storage; and silicon photovoltaics.
An initial analysis of the data from the Polar Orbiting Geophysical (POGS) Satellite
NASA Technical Reports Server (NTRS)
Langel, R. A.; Sabaka, T. J.; Baldwin, R. T.
1991-01-01
The Polar Orbiting Geophysical Satellite (POGS) was launched in 1990 to measure the geomagnetic field. POGS data from selected magnetically quiet days was chosen, quality checked and deleted where thought to be erroneous. A time and position correction was applied. The resulting data was fit to a degree 13 spherical harmonic model. Evaluation of the quality of the data indicates that it is sufficient for definition of the low degree (approximately less than 8) portion of the geomagnetic field. Further correction of the data time and position may improve this quality.
Initial effects of nuclear weapon x-radiation on the LAMPSHADE orbital debris satellite shield
Smith, M.S.; Santoro, R.T.
1989-09-01
One-dimensional thermal-hydrodynamic calculations have been carried out to estimate the response of the lead bumper plate and tantalum liquidation screen of the LAMPSHADE orbital debris satellite shield. The mass loss fraction in the solid, liquid, and vapor phases as a function of time after irradiation for several typical incident x-ray spectra fluences were calculated using the PUFF-TFT code. The material losses did not exceed 2% and fracture and spallation were confined to the surface region with no apparent reduction in the performance of these components against incident debris. 4 refs., 5 figs., 3 tabs.
NASA Technical Reports Server (NTRS)
Goossens, S.; Matsumoto, K.; Noda, H.; Araki, H.; Rowlands, D. D.; Lemoine, F. G.
2011-01-01
The SELENE mission, consisting of three separate satellites that use different terrestrial-based tracking systems, presents a unique opportunity to evaluate the contribution of these tracking systems to orbit determination precision. The tracking data consist of four-way Doppler between the main orbiter and one of the two sub-satellites while the former is over the far side, and of same-beam differential VLBI tracking between the two sub-satellites. Laser altimeter data are also used for orbit determination. The contribution to orbit precision of these different data types is investigated through orbit overlap analysis. It is shown that using four-way and VLBI data improves orbit consistency for all satellites involved by reducing peak values in orbit overlap differences that exist when only standard two-way Doppler and range data are used. Including laser altimeter data improves the orbit precision of the SELENE main satellite further, resulting in very smooth total orbit errors at an average level of 18m. The multi-satellite data have also resulted in improved lunar gravity field models, which are assessed through orbit overlap analysis using Lunar Prospector tracking data. Improvements over a pre-SELENE model are shown to be mostly in the along-track and cross-track directions. Orbit overlap differences are at a level between 13 and 21 m with the SELENE models, depending on whether l-day data overlaps or I-day predictions are used.
Flight dynamics facility operational orbit determination support for the ocean topography experiment
NASA Technical Reports Server (NTRS)
Bolvin, D. T.; Schanzle, A. F.; Samii, M. V.; Doll, C. E.
1991-01-01
The Ocean Topography Experiment (TOPEX/POSEIDON) mission is designed to determine the topography of the Earth's sea surface across a 3 yr period, beginning with launch in June 1992. The Goddard Space Flight Center Dynamics Facility has the capability to operationally receive and process Tracking and Data Relay Satellite System (TDRSS) tracking data. Because these data will be used to support orbit determination (OD) aspects of the TOPEX mission, the Dynamics Facility was designated to perform TOPEX operational OD. The scientific data require stringent OD accuracy in navigating the TOPEX spacecraft. The OD accuracy requirements fall into two categories: (1) on orbit free flight; and (2) maneuver. The maneuver OD accuracy requirements are of two types; premaneuver planning and postmaneuver evaluation. Analysis using the Orbit Determination Error Analysis System (ODEAS) covariance software has shown that, during the first postlaunch mission phase of the TOPEX mission, some postmaneuver evaluation OD accuracy requirements cannot be met. ODEAS results also show that the most difficult requirements to meet are those that determine the change in the components of velocity for postmaneuver evaluation.
Federal Register 2010, 2011, 2012, 2013, 2014
2013-08-07
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NASA Astrophysics Data System (ADS)
Lei, YANG; Caifa, GUO; Zhengxu, DAI; Xiaoyong, LI; Shaolin, WANG
2016-02-01
The space tracking ship is a moving platform in the TT&C network. The orbit determination precision of the ship plays a key role in the TT&C mission. Based on the measuring data obtained by the ship-borne equipments, the paper presents the mathematic models of the complicated error from the space tracking ship, which can separate the random error and the correction residual error with secondary low frequency from the complicated error. An error simulation algorithm is proposed to analyze the orbit determination precision based on the two set of the different equipments. With this algorithm, a group of complicated error can be simulated from a measured sample. The simulated error groups can meet the requirements of sufficient complicated error for the equipment tests before the mission execution, which is helpful to the practical application.
Atmospheric drag model for Cassini orbit determination during low altitude Titan flybys
NASA Technical Reports Server (NTRS)
Pelletier, F. J.; Antreasian, P. G.; Bordi, J. J.; Criddle, K. E.; Ionasescu, R.; Jacobson, R. A.; Mackenzie, R. A.; Parcher, D. W.; Stauch, J. R.
2006-01-01
On April 16, 2005, the Cassini spacecraft performed its lowest altitude flyby of Titan to date, the Titan-5 flyby, flying 1027 km above the surface of Titan. This document discusses the development of a Titan atmospheric drag model for the purpose of the orbit determination of Cassini. Results will be presented for the Titan A flyby, the Titan-5 flyby as well as the most recent low altitude Titan flyby, Titan-7. Different solutions will be compared against OD performance in terms of the flyby B-plane parameters, spacecraft thrusting activity and drag estimates. These low altitude Titan flybys were an excellent opportunity to observe the effect of Titan's atmospheric drag on the orbit determination solution and results show that the drag was successfully modeled to provide accurate flyby solutions.
Comparison of Sigma-Point and Extended Kalman Filters on a Realistic Orbit Determination Scenario
NASA Technical Reports Server (NTRS)
Gaebler, John; Hur-Diaz. Sun; Carpenter, Russell
2010-01-01
Sigma-point filters have received a lot of attention in recent years as a better alternative to extended Kalman filters for highly nonlinear problems. In this paper, we compare the performance of the additive divided difference sigma-point filter to the extended Kalman filter when applied to orbit determination of a realistic operational scenario based on the Interstellar Boundary Explorer mission. For the scenario studied, both filters provided equivalent results. The performance of each is discussed in detail.
TOPEX orbit determination by solving gravity parameters with multiple arc data
NASA Technical Reports Server (NTRS)
Wu, J.-T.
1986-01-01
Multiple arc data from repeated ground track are combined to reduce the error due to gravity field uncertainty in the determination of TOPEX orbit. The TOPEX dynamics is modeled with relatively few gravity parameters to account for the effect of the local gravity field. The gravity parameters are common to all arcs. The estimation algorithm uses the Householder transformation to combine multiple arc data and solve for the gravity parameters. The earth gravity field can be recovered with very modest amount of calculation.
Precise orbit determination of Multi-GNSS constellation including GPS GLONASS BDS and GALIEO
NASA Astrophysics Data System (ADS)
Dai, Xiaolei
2014-05-01
In addition to the existing American global positioning system (GPS) and the Russian global navigation satellite system (GLONASS), the new generation of GNSS is emerging and developing, such as the Chinese BeiDou satellite navigation system (BDS) and the European GALILEO system. Multi-constellation is expected to contribute to more accurate and reliable positioning and navigation service. However, the application of multi-constellation challenges the traditional precise orbit determination (POD) strategy that was designed usually for single constellation. In this contribution, we exploit a more rigorous multi-constellation POD strategy for the ongoing IGS multi-GNSS experiment (MGEX) where the common parameters are identical for each system, and the frequency- and system-specified parameters are employed to account for the inter-frequency and inter-system biases. Since the authorized BDS attitude model is not yet released, different BDS attitude model are implemented and their impact on orbit accuracy are studied. The proposed POD strategy was implemented in the PANDA (Position and Navigation Data Analyst) software and can process observations from GPS, GLONASS, BDS and GALILEO together. The strategy is evaluated with the multi-constellation observations from about 90 MGEX stations and BDS observations from the BeiDou experimental tracking network (BETN) of Wuhan University (WHU). Of all the MGEX stations, 28 stations record BDS observation, and about 80 stations record GALILEO observations. All these data were processed together in our software, resulting in the multi-constellation POD solutions. We assessed the orbit accuracy for GPS and GLONASS by comparing our solutions with the IGS final orbit, and for BDS and GALILEO by overlapping our daily orbit solution. The stability of inter-frequency bias of GLONASS and inter-system biases w.r.t. GPS for GLONASS, BDS and GALILEO were investigated. At last, we carried out precise point positioning (PPP) using the multi-constellation POD orbit and clock products, and analyzed the contribution of these POD products to PPP. Keywords: Multi-GNSS, Precise Orbit Determination, Inter-frequency bias, Inter-system bias, Precise Point Positioning
Orbit determination across unknown maneuvers using the essential Thrust-Fourier-Coefficients
NASA Astrophysics Data System (ADS)
Ko, Hyun Chul; Scheeres, Daniel J.
2016-01-01
Any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver involving Thrust-Fourier-Coefficients (TFCs). With a selected TFC set as a basis, a thrust acceleration can be constructed to interpolate two unconnected states across an unknown maneuver. This representation technique with TFCs enables us to facilitate the analytical propagation of uncertainties of the satellite state. This approach allows for the usage of existing pre-maneuver orbit estimation to compute the orbit solution after the unknown maneuver. In this paper, we applied this approach to orbit determination (OD) problems across unknown maneuvers by appending different combinations of TFCs to the state vector in the batch filter. The aim is to investigate how different maneuver representations with different TFC sets affect the OD solution across unknown maneuvers. Simulation results show that each TFC set provides different representations of the unknown perturbing acceleration, which yields varying magnitudes of delta velocity for a given maneuver. However, OD solutions across unknown maneuvers using different TFC sets display equivalent performance over the post-maneuver arc as long as those TFC sets are capable of generating the apparent secular motion caused by a given unknown maneuver.
NASA Technical Reports Server (NTRS)
Peters, Palmer N.; Gregory, John C.
1991-01-01
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.
A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver
Wang, Fuhong; Gong, Xuewen; Sang, Jizhang; Zhang, Xiaohong
2015-01-01
Satellite remote sensing systems require accurate, autonomous and real-time orbit determinations (RTOD) for geo-referencing. Onboard Global Positioning System (GPS) has widely been used to undertake such tasks. In this paper, a novel RTOD method achieving decimeter precision using GPS carrier phases, required by Chinaâ€™s HY2A and ZY3 missions, is presented. A key to the algorithm success is the introduction of a new parameter, termed pseudo-ambiguity. This parameter combines the phase ambiguity, the orbit, and clock offset errors of the GPS broadcast ephemeris together to absorb a large part of the combined error. Based on the analysis of the characteristics of the orbit and clock offset errors, the pseudo-ambiguity can be modeled as a random walk, and estimated in an extended Kalman filter. Experiments of processing real data from HY2A and ZY3, simulating onboard operational scenarios of these two missions, are performed using the developed software SATODS. Results have demonstrated that the position and velocity accuracy (3D RMS) of 0.2â€“0.4 m and 0.2â€“0.4 mm/s, respectively, are achieved using dual-frequency carrier phases for HY2A, and slightly worse results for ZY3. These results show it is feasible to obtain orbit accuracy at decimeter level of 3â€“5 dm for position and 0.3â€“0.5 mm/s for velocity with this RTOD method. PMID:26690149
A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver.
Wang, Fuhong; Gong, Xuewen; Sang, Jizhang; Zhang, Xiaohong
2015-01-01
Satellite remote sensing systems require accurate, autonomous and real-time orbit determinations (RTOD) for geo-referencing. Onboard Global Positioning System (GPS) has widely been used to undertake such tasks. In this paper, a novel RTOD method achieving decimeter precision using GPS carrier phases, required by China's HY2A and ZY3 missions, is presented. A key to the algorithm success is the introduction of a new parameter, termed pseudo-ambiguity. This parameter combines the phase ambiguity, the orbit, and clock offset errors of the GPS broadcast ephemeris together to absorb a large part of the combined error. Based on the analysis of the characteristics of the orbit and clock offset errors, the pseudo-ambiguity can be modeled as a random walk, and estimated in an extended Kalman filter. Experiments of processing real data from HY2A and ZY3, simulating onboard operational scenarios of these two missions, are performed using the developed software SATODS. Results have demonstrated that the position and velocity accuracy (3D RMS) of 0.2-0.4 m and 0.2-0.4 mm/s, respectively, are achieved using dual-frequency carrier phases for HY2A, and slightly worse results for ZY3. These results show it is feasible to obtain orbit accuracy at decimeter level of 3-5 dm for position and 0.3-0.5 mm/s for velocity with this RTOD method. PMID:26690149
Relative orbit determination for satellite formation flying based on quantum ranging
NASA Astrophysics Data System (ADS)
Shen, Yanghe; Xu, Luping; Zhang, Hua; Chen, Shanshan; Song, Shibin
2015-08-01
Relative orbit determination is widely used in the field of autonomously controlled satellite formation flying (SFF). Currently, some traditional techniques cannot meet the strict requirement of the accuracy of relative orbit determination for certain space missions. Thus, the primary purpose of this study is to design some special type of sensor to increase the accuracy of the distance measurement, which can eventually lead to an improvement in the accuracy of relative orbit determination for SFF. Two types of quantum sensors are proposed, based on the double-points quantum ranging (DPQR) and the triangle quantum ranging (TQR) schemes that utilize the second-order correlation between the entangled photons. Simulation result shows that the ranging accuracy of the TQR-type sensor is more precise than that of the DPQR-type one. Additionally, the unscented Kalman filter (UKF) is used to estimate the relative state of the SFF, which uses the TQR-type sensor as the measurement model compared with a traditional sensor. The simulation results show that the quantum sensor is superior to the traditional one and their estimation errors of the position and velocity remain within 1 cm and 1 mm/s, respectively, at a relative distance of 1 km between the chief and deputy satellites.
Determination of spin and orbital magnetization in the ferromagnetic superconductor UCoGe
NASA Astrophysics Data System (ADS)
Butchers, M. W.; Duffy, J. A.; Taylor, J. W.; Giblin, S. R.; Dugdale, S. B.; Stock, C.; Tobash, P. H.; Bauer, E. D.; Paulsen, C.
2015-09-01
The magnetism in the ferromagnetic superconductor UCoGe has been studied using a combination of magnetic Compton scattering, bulk magnetization, x-ray magnetic circular dichroism, and electronic structure calculations, in order to determine the spin and orbital moments. The experimentally observed total spin moment Ms was found to be -0.24 ±0.05 ?B at 5 T. By comparison with the total moment of 0.16 ±0.01 ?B , the orbital moment Ml was determined to be 0.40 ±0.05 ?B . The U and Co spin moments were determined to be antiparallel. We find that the U 5 f electrons carry a spin moment of Us?-0.30 ?B and that there is a Co spin moment of Cos?0.06 ?B induced via hybridization. The ratio Ul/Us , of -1.3 ±0.3 , shows the U moment to be itinerant. In order to ensure an accurate description of the properties of 5 f systems, and to provide a critical test of the theoretical approaches, it is clearly necessary to obtain experimental data for both the spin and orbital moments, rather than just the total magnetic moment. This can be achieved simply by measuring the spin moment with magnetic Compton scattering and comparing this to the total moment from bulk magnetization.
DETERMINING THE INITIAL HELIUM ABUNDANCE OF THE SUN
Serenelli, Aldo M.; Basu, Sarbani
2010-08-10
We determine the dependence of the initial helium abundance and the present-day helium abundance in the convective envelope of solar models (Y {sub ini} and Y {sub surf}, respectively) on the parameters that are used to construct the models. We do so by using reference standard solar models (SSMs) to compute the power-law coefficients of the dependence of Y {sub ini} and Y {sub surf} on the input parameters. We use these dependencies to determine the correlation between Y {sub ini} and Y {sub surf} and use this correlation to eliminate uncertainties in Y {sub ini} from all solar model input parameters except the microscopic diffusion rate. We find an expression for Y {sub ini} that depends only on Y {sub surf} and the diffusion rate. By adopting the helioseismic determination of solar surface helium abundance, Y {sup surf} {sub sun} = 0.2485 {+-} 0.0035, and an uncertainty of 20% for the diffusion rate, we find that the initial solar helium abundance, Y {sup ini} {sub sun}, is 0.278 {+-} 0.006 independently of the reference SSMs (and particularly on the adopted solar abundances) used in the derivation of the correlation between Y {sub ini} and Y {sub surf}. When non-SSMs with extra mixing are used, then we derive Y {sup ini} {sub sun} = 0.273 {+-} 0.006. In both cases, the derived Y {sup ini} {sub sun} value is higher than that directly derived from solar model calibrations when the low-metallicity solar abundances (e.g., by Asplund et al.) are adopted in the models.
Lunar Reconnaissance Orbiter Camera Narrow Angle Cameras: Laboratory and Initial Flight Calibration
NASA Astrophysics Data System (ADS)
Humm, D. C.; Tschimmel, M.; Denevi, B. W.; Lawrence, S.; Mahanti, P.; Tran, T. N.; Thomas, P. C.; Eliason, E.; Robinson, M. S.
2009-12-01
The Lunar Reconnaissance Orbiter Camera (LROC) has two identical Narrow Angle Cameras (NACs). Each NAC is a monochrome pushbroom scanner, providing images with a pixel scale of 50 cm from a 50-km orbit. A single NAC image has a swath width of 2.5 km and a length of up to 26 km. The NACs are mounted to acquire side-by-side imaging for a combined swath width of 5 km. The NAC is designed to fully characterize future human and robotic landing sites in terms of scientific and resource merit, trafficability, and hazards. The North and South poles will be mapped at 1-meter-scale poleward of 85.5 degrees latitude. Stereo coverage is achieved by pointing the NACs off-nadir, which requires planning in advance. Read noise is 91 and 93 e- and the full well capacity is 334,000 and 352,000 e- for NAC-L and NAC-R respectively. Signal-to-noise ranges from 42 for low-reflectance material with 70 degree illumination to 230 for high-reflectance material with 0 degree illumination. Longer exposure times and 2x binning are available to further increase signal-to-noise with loss of spatial resolution. Lossy data compression from 12 bits to 8 bits uses a companding table selected from a set optimized for different signal levels. A model of focal plane temperatures based on flight data is used to command dark levels for individual images, optimizing the performance of the companding tables and providing good matching of the NAC-L and NAC-R images even before calibration. The preliminary NAC calibration pipeline includes a correction for nonlinearity at low signal levels with an offset applied for DN>600 and a logistic function for DN<600. Flight images taken on the limb of the Moon provide a measure of stray light performance. Averages over many lines of images provide a measure of flat field performance in flight. These are comparable with laboratory data taken with a diffusely reflecting uniform panel.
Determination of the force transmitted by an ion thruster plasma plume to an orbital object
NASA Astrophysics Data System (ADS)
Alpatov, A.; Cichocki, F.; Fokov, A.; Khoroshylov, S.; Merino, M.; Zakrzhevskii, A.
2016-02-01
An approach to determine the force transmitted by the plasma plume of an ion thruster to an orbital object immersed in it using its central projection on a selected plane is proposed. A photo camera is used to obtain the image of the object central projection. The algorithms for the calculation of the transmission of momentum by the impacting ion beam are developed including the determination of the object contour and the correction of the error due to a camera offset from the ion beam axis, and the computation of the fraction of the ion beam that impinges on the object surface.
NASA Astrophysics Data System (ADS)
Jorgensen, Kira; Africano, John L.; Stansbery, Eugene G.; Kervin, Paul W.; Hamada, Kris M.; Sydney, Paul F.
2001-12-01
The purpose of this research is to improve the knowledge of the physical properties of orbital debris, specifically the material type. Combining the use of the fast-tracking United States Air Force Research Laboratory (AFRL) telescopes with a common astronomical technique, spectroscopy, and NASA resources was a natural step toward determining the material type of orbiting objects remotely. Currently operating at the AFRL Maui Optical Site (AMOS) is a 1.6-meter telescope designed to track fast moving objects like those found in lower Earth orbit (LEO). Using the spectral range of 0.4 - 0.9 microns (4000 - 9000 angstroms), researchers can separate materials into classification ranges. Within the above range, aluminum, paints, plastics, and other metals have different absorption features as well as slopes in their respective spectra. The spectrograph used on this telescope yields a three-angstrom resolution; large enough to see smaller features mentioned and thus determine the material type of the object. The results of the NASA AMOS Spectral Study (NASS) are presented herein.
Precise Orbit Determination of LEO Satellite Using Dual-Frequency GPS Data
NASA Astrophysics Data System (ADS)
Hwang, Yoola; Lee, Byoung-Sun; Kim, Jaehoon; Yoon, Jae-Cheol
2009-06-01
KOrea Multi-purpose SATellite (KOMPSAT)-5 will be launched at 550km altitude in 2010. Accurate satellite position (20 cm) and velocity (0.03 cm/s) are required to treat highly precise Synthetic Aperture Radar (SAR) image processing. Ionosphere delay was eliminated using dual frequency GPS data and double differenced GPS measurement removed common clock errors of both GPS satellites and receiver. SAC-C carrier phase data with 0.1 Hz sampling rate was used to achieve precise orbit determination (POD) with ETRI GNSS Precise Orbit Determination (EGPOD) software, which was developed by ETRI. Dynamic model approach was used and satellite's position, velocity, and the coefficients of solar radiation pressure and drag were adjusted once per arc using Batch Least Square Estimator (BLSE) filter. Empirical accelerations for sinusoidal radial, along-track, and cross track terms were also estimated once per revolution for unmodeled dynamics. Additionally piece-wise constant acceleration for cross-track direction was estimated once per arc. The performance of POD was validated by comparing with JPL's Precise Orbit Ephemeris (POE).
NASA Technical Reports Server (NTRS)
Stephenson, Frank W., Jr.
1988-01-01
The NASA Earth-to-Orbit (ETO) Propulsion Technology Program is dedicated to advancing rocket engine technologies for the development of fully reusable engine systems that will enable space transportation systems to achieve low cost, routine access to space. The program addresses technology advancements in the areas of engine life extension/prediction, performance enhancements, reduced ground operations costs, and in-flight fault tolerant engine operations. The primary objective is to acquire increased knowledge and understanding of rocket engine chemical and physical processes in order to evolve more realistic analytical simulations of engine internal environments, to derive more accurate predictions of steady and unsteady loads, and using improved structural analyses, to more accurately predict component life and performance, and finally to identify and verify more durable advanced design concepts. In addition, efforts were focused on engine diagnostic needs and advances that would allow integrated health monitoring systems to be developed for enhanced maintainability, automated servicing, inspection, and checkout, and ultimately, in-flight fault tolerant engine operations.
NASA Technical Reports Server (NTRS)
Kliore, A. J.; Patel, I. R.; Nagy, A. F.; Cravens, T. E.; Gombosi, T. I.
1979-01-01
Results of radio occultation measurements of electron density profiles of the nightside ionosphere of Venus at solar zenith angles from 90 to 164 deg, obtained from the Pioneer Venus Orbiter, are reported. Data were derived from closed-loop S- and X-band signals received by the Deep Space Network upon ionospheric entry and exit of the spacecraft. Nightside electron density profiles are found to be rather uniform in the solar zenith angle range of from 95 to 107 deg, with peak electron densities ranging from 23,000 to 40,000/cu cm, while between 110 and 164 deg, profiles exhibit a high degree of variability and peak electron densities vary from 7,600 to 31,800/cu cm. A possible mechanism for the maintenance of the nightside Venus ionosphere during the long Venus night, which is consistent with the observed spatial and temporal variability of deep ionospheric electron density profiles, is proposed to be impact ionization by precipitating particles, although transport processes from the dayside may also be important.
GPS-based orbit determination and point positioning under selective availability
NASA Astrophysics Data System (ADS)
Bar-Sever, Yoaz E.; Yunck, Thomas P.; Wu, Sien-Chong
Selective availability (SA) degrades the positioning accuracy for nondifferential users of the GPS Standard Positioning Service (SPS). The often quoted SPS accuracy available under normal conditions is 100 m 2DRMS. In the absence of more specific information, many prospective SPS users adopt the 100 m value in their planning, which exaggerates the error in many cases. SA error is examined for point positioning and dynamic orbit determination for an orbiting user. To minimize SA error, nondifferential users have several options: expand their field of view; observe as many GPS satellites as possible; smooth the error over time; and employ precise GPS ephemerides computed independently, as by NASA and the NGS, rather than the broadcast ephemeris. Simulations show that 3D point position error can be kept to 30 m, and this can be smoothed to 3 m in a few hours.
Precise Orbit Determination for LEO Spacecraft Using GNSS Tracking Data from Multiple Antennas
NASA Technical Reports Server (NTRS)
Kuang, Da; Bertiger, William; Desai, Shailen; Haines, Bruce
2010-01-01
To support various applications, certain Earth-orbiting spacecrafts (e.g., SRTM, COSMIC) use multiple GNSS antennas to provide tracking data for precise orbit determination (POD). POD using GNSS tracking data from multiple antennas poses some special technical issues compared to the typical single-antenna approach. In this paper, we investigate some of these issues using both real and simulated data. Recommendations are provided for POD with multiple GNSS antennas and for antenna configuration design. The observability of satellite position with multiple antennas data is compared against single antenna case. The impact of differential clock (line biases) and line-of-sight (up, along-track, and cross-track) on kinematic and reduced-dynamic POD is evaluated. The accuracy of monitoring the stability of the spacecraft structure by simultaneously performing POD of the spacecraft and relative positioning of the multiple antennas is also investigated.
GPS-based orbit determination and point positioning under selective availability
NASA Technical Reports Server (NTRS)
Bar-Sever, Yoaz E.; Yunck, Thomas P.; Wu, Sien-Chong
1990-01-01
Selective availability (SA) degrades the positioning accuracy for nondifferential users of the GPS Standard Positioning Service (SPS). The often quoted SPS accuracy available under normal conditions is 100 m 2DRMS. In the absence of more specific information, many prospective SPS users adopt the 100 m value in their planning, which exaggerates the error in many cases. SA error is examined for point positioning and dynamic orbit determination for an orbiting user. To minimize SA error, nondifferential users have several options: expand their field of view; observe as many GPS satellites as possible; smooth the error over time; and employ precise GPS ephemerides computed independently, as by NASA and the NGS, rather than the broadcast ephemeris. Simulations show that 3D point position error can be kept to 30 m, and this can be smoothed to 3 m in a few hours.
NASA Astrophysics Data System (ADS)
Shoji, Mitsuo; Yoshioka, Yasunori; Yamaguchi, Kizashi
2014-07-01
A novel procedure to generate initial broken-symmetry solutions is proposed. Conventional methods for the initial broken-symmetry solutions are the MO alter, HOMO-LUMO mixing and fragment methods. These procedures, however, are quite complex. Our new approach is efficient, automatic and highly practical especially for large QM systems. This approach, called the LNO method, is applied to the following four typical open-shell systems: H2, dicarbene and two iron-sulfur clusters of Rieske-type [2Fe-2S] and [4Fe-4S]. The performance and the efficiency as an electronic structural analysis are discussed. The LNO method will be applicable for general systems in the complicated broken symmetry states.
NASA Technical Reports Server (NTRS)
Marr, Greg C.
2003-01-01
The Triana spacecraft was designed to be launched by the Space Shuttle. The nominal Triana mission orbit will be a Sun-Earth L1 libration point orbit. Using the NASA Goddard Space Flight Center's Orbit Determination Error Analysis System (ODEAS), orbit determination (OD) error analysis results are presented for all phases of the Triana mission from the first correction maneuver through approximately launch plus 6 months. Results are also presented for the science data collection phase of the Fourier Kelvin Stellar Interferometer Sun-Earth L2 libration point mission concept with momentum unloading thrust perturbations during the tracking arc. The Triana analysis includes extensive analysis of an initial short arc orbit determination solution and results using both Deep Space Network (DSN) and commercial Universal Space Network (USN) statistics. These results could be utilized in support of future Sun-Earth libration point missions.
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2010-12-01
A new method is suggested for computing the initial orbit of a small celestial body from its three or more pairs of angular measurements at three times. The method is based on using the approach that we previously developed for constructing the intermediate orbit from minimal number of observations. This intermediate orbit allows for most of the perturbations in the motion of the body under study. The method proposed uses the Herget's algorithmic scheme that makes it possible to involve additional observations as well. The methodical error of orbit computation by the proposed method is two orders smaller than the corresponding error of the Herget's approach based on the construction of the unperturbed Keplerian orbit. The new method is especially efficient if applied to high-accuracy observational data covering short orbital arcs.
Evaluation of LANDSAT-D Orbit Determination Using a Filter/Smoother (PREFER)
NASA Technical Reports Server (NTRS)
Gibbs, B. P.
1982-01-01
Simulated range and range rate data for five tracking stations were first generated using batch least squares orbit determination (GTDS). Then GTDS was used (in the differential correction mode) to produce a nominal trajectory which was input to PREFER. The GTDS differential correction (DC) run was made using models which differed from those used to produce the simulated data. These model differences were chosen to be fairly realistic approximations to the errors in the models actually used for operational orbit determination. Several different simulation runs were made with different types of model errors in order to determine the sensitivity to these errors. The nominal trajectory and the simulated measurement data were input to PREFER to produce a smoothed ephemeris file. Numerous runs of PREFER were made in which parameters describing the statistics of the model errors were varied. The likelihood function computed by the Kalman filter determined the ""best'' choice of input parameters. There was strong negative correlation between the likelihood function and the errors in the smoothed ephemeris.
Modeling of Non-Gravitational Forces for Precise and Accurate Orbit Determination
NASA Astrophysics Data System (ADS)
Hackel, Stefan; Gisinger, Christoph; Steigenberger, Peter; Balss, Ulrich; Montenbruck, Oliver; Eineder, Michael
2014-05-01
Remote sensing satellites support a broad range of scientific and commercial applications. The two radar imaging satellites TerraSAR-X and TanDEM-X provide spaceborne Synthetic Aperture Radar (SAR) and interferometric SAR data with a very high accuracy. The precise reconstruction of the satellite's trajectory is based on the Global Positioning System (GPS) measurements from a geodetic-grade dual-frequency Integrated Geodetic and Occultation Receiver (IGOR) onboard the spacecraft. The increasing demand for precise radar products relies on validation methods, which require precise and accurate orbit products. An analysis of the orbit quality by means of internal and external validation methods on long and short timescales shows systematics, which reflect deficits in the employed force models. Following the proper analysis of this deficits, possible solution strategies are highlighted in the presentation. The employed Reduced Dynamic Orbit Determination (RDOD) approach utilizes models for gravitational and non-gravitational forces. A detailed satellite macro model is introduced to describe the geometry and the optical surface properties of the satellite. Two major non-gravitational forces are the direct and the indirect Solar Radiation Pressure (SRP). The satellite TerraSAR-X flies on a dusk-dawn orbit with an altitude of approximately 510 km above ground. Due to this constellation, the Sun almost constantly illuminates the satellite, which causes strong across-track accelerations on the plane rectangular to the solar rays. The indirect effect of the solar radiation is called Earth Radiation Pressure (ERP). This force depends on the sunlight, which is reflected by the illuminated Earth surface (visible spectra) and the emission of the Earth body in the infrared spectra. Both components of ERP require Earth models to describe the optical properties of the Earth surface. Therefore, the influence of different Earth models on the orbit quality is assessed. The scope of the presentation is a detailed analysis of the orbit improvements due to sophisticated non-gravitational force and satellite macro models for the satellite TerraSAR-X.
Representation of Probability Density Functions from Orbit Determination using the Particle Filter
NASA Technical Reports Server (NTRS)
Mashiku, Alinda K.; Garrison, James; Carpenter, J. Russell
2012-01-01
Statistical orbit determination enables us to obtain estimates of the state and the statistical information of its region of uncertainty. In order to obtain an accurate representation of the probability density function (PDF) that incorporates higher order statistical information, we propose the use of nonlinear estimation methods such as the Particle Filter. The Particle Filter (PF) is capable of providing a PDF representation of the state estimates whose accuracy is dependent on the number of particles or samples used. For this method to be applicable to real case scenarios, we need a way of accurately representing the PDF in a compressed manner with little information loss. Hence we propose using the Independent Component Analysis (ICA) as a non-Gaussian dimensional reduction method that is capable of maintaining higher order statistical information obtained using the PF. Methods such as the Principal Component Analysis (PCA) are based on utilizing up to second order statistics, hence will not suffice in maintaining maximum information content. Both the PCA and the ICA are applied to two scenarios that involve a highly eccentric orbit with a lower apriori uncertainty covariance and a less eccentric orbit with a higher a priori uncertainty covariance, to illustrate the capability of the ICA in relation to the PCA.
NASA Technical Reports Server (NTRS)
Lyons, Frankel
2013-01-01
A new orbital debris environment model (ORDEM 3.0) defines the density distribution of the debris environment in terms of the fraction of debris that are low-density (plastic), medium-density (aluminum) or high-density (steel) particles. This hypervelocity impact (HVI) program focused on assessing ballistic limits (BLs) for steel projectiles impacting the enhanced Soyuz Orbital Module (OM) micrometeoroid and orbital debris (MMOD) shield configuration. The ballistic limit was defined as the projectile size on the threshold of failure of the OM pressure shell as a function of impact speeds and angle. The enhanced OM shield configuration was first introduced with Soyuz 30S (launched in May 2012) to improve the MMOD protection of Soyuz vehicles docked to the International Space Station (ISS). This test program provides HVI data on U.S. materials similar in composition and density to the Russian materials for the enhanced Soyuz OM shield configuration of the vehicle. Data from this test program was used to update ballistic limit equations used in Soyuz OM penetration risk assessments. The objective of this hypervelocity impact test program was to determine the ballistic limit particle size for 440C stainless steel spherical projectiles on the Soyuz OM shielding at several impact conditions (velocity and angle combinations). This test report was prepared by NASA-JSC/ HVIT, upon completion of tests.
The challenge of precise orbit determination for STSAT-2C using extremely sparse SLR data
NASA Astrophysics Data System (ADS)
Kim, Young-Rok; Park, Eunseo; Kucharski, Daniel; Lim, Hyung-Chul; Kim, Byoungsoo
2016-03-01
The Science and Technology Satellite (STSAT)-2C is the first Korean satellite equipped with a laser retro-reflector array for satellite laser ranging (SLR). SLR is the only on-board tracking source for precise orbit determination (POD) of STSAT-2C. However, POD for the STSAT-2C is a challenging issue, as the laser measurements of the satellite are extremely sparse, largely due to the inaccurate two-line element (TLE)-based orbit predictions used by the SLR tracking stations. In this study, POD for the STSAT-2C using extremely sparse SLR data is successfully implemented, and new laser-based orbit predictions are obtained. The NASA/GSFC GEODYN II software and seven-day arcs are used for the SLR data processing of two years of normal points from March 2013 to May 2015. To compensate for the extremely sparse laser tracking, the number of estimation parameters are minimized, and only the atmospheric drag coefficients are estimated with various intervals. The POD results show that the weighted root mean square (RMS) post-fit residuals are less than 10 m, and the 3D day boundaries vary from 30 m to 3 km. The average four-day orbit overlaps are less than 20/330/20 m for the radial/along-track/cross-track components. The quality of the new laser-based prediction is verified by SLR observations, and the SLR residuals show better results than those of previous TLE-based predictions. This study demonstrates that POD for the STSAT-2C can be successfully achieved against extreme sparseness of SLR data, and the results can deliver more accurate predictions.
Determination of the Orbit of the Planetary Companion to the Metal-Rich Star HD 45350
NASA Astrophysics Data System (ADS)
Endl, Michael; Cochran, William D.; Wittenmyer, Robert A.; Hatzes, Artie P.
2006-06-01
We present precise radial velocity data for the metal-rich star HD 45350 collected with the Harlan J. Smith (HJS) 2.7 m telescope and the Hobby-Eberly Telescope (HET) at McDonald Observatory. This star was noticed by us as a candidate for having a giant planetary companion in a highly eccentric orbit, but the lack of data close to periastron left the amplitude and thus the mass of the planet poorly constrained. Marcy et al. (2005) announced the presence of the planet based on their Keck HIRES data, but those authors also cautioned that the remaining uncertainties in the orbital solution might be large due to insufficient data near periastron passage. In order to close this phase gap we exploited the flexible queue-scheduled observing mode of the HET to obtain intensive coverage of the most recent periastron passage of the planet. In combination with the long-term data from the HJS 2.7 m telescope we determine a Keplerian orbital solution for this system with a period of 962 days, an eccentricity of e=0.76, and a velocity semiamplitude K of 57.4 m s-1. The planet has a minimum mass of msini=1.82MJ+/-0.14MJ and an orbital semimajor axis of a=1.92+/-0.07 AU. Based on observations obtained with the Harlan J. Smith Telescope and the Hobby-Eberly Telescope, which is a joint project of the University of Texas at Austin, Pennsylvania State University, Stanford University, Ludwig-Maximilians-Universität München, and Georg-August-Universität Göttingen.
Relativistic Orbit Determination for The Lisa Mission: a Numerical Versus an Analytical Approach
NASA Astrophysics Data System (ADS)
Pireaux, Sophie M. V.; Chauvineau, B.
2009-05-01
The LISA mission is an interferometer, formed by three spacecraft, that aims at the detection of gravitational waves in the [10-4, 10-1] Hz frequency band. Before the present work, only CLASSICAL ephemerides for LISA satellites were available. Hence, relativistic effects in LISA orbit determination needed to be considered and quantified. We consider here the motion of LISA satellites in the gravitational field of a spherical non-rotating Sun, without planets. The Relativistic Motion Integrator (RMI) consists in integrating numerically the exact equations of motion (geodesics), for a given metric. The RMI approach can be applied to compute either planetocentric or barycentric orbits for different space missions. As an application, we consider a relativistic metric which corresponds to a gravitational field at first post-Newtonian order. Indeed, LISA is a relevant example to use RMI together with a BCRS (Barycentric Coordinate Reference System) metric, as recommended by the IAU (International Astronomical Union). To validate RMI's results for LISA, we used an analytical development (up to first order in the eccentricity e of the orbit and up to first order in GM/c2, where G is Newton's constant, M, the solar mass and c the speed of light in vacuum). We show that RMI's results agree with this analytical development within e2 GM/c2. We show that a numerical classical model for LISA orbits in the gravitational field of a non-rotating spherical Sun without planets can be wrong, with respect to the relativistic version of the same model, by as much as about ten kilometers in radial distance during a year and up to about 60 kilometer in along track distance after a year... with consequences on estimated photon flight times. "Relativistic versus Newtonian orbitography: RMI software, illustration with the LISA mission", S. Pireaux, B. Chauvineau, arXiv:0801.3627v1(gr-qc)
Orbit Determination of Chang'e-3 and Positioning of the Lander and the Rover
NASA Astrophysics Data System (ADS)
Huang, Y.; Chang, S.; Li, P.; Hu, X.
2014-12-01
The Chang'E-3 (CE-3) lunar probe of China was launched on 2 December 2013. After about 112 h of flight, it was captured by the Moon on 6 December, and entered a polar, near circular lunar orbit with an altitude of approximately 100 km. The probe's flight on 100 km*100 km and 100 km*15 km orbit lasted about 4 days respectively, then the probe soft landed on the east of Sinus Iridum area at 13:11 UTC on 14 December successfully. Results on precision orbit determination and positioning of the lander and the rover are presented here. We describe the data, modeling and methods used to achieve position knowledge. In addition to the radiometric X-band range and Doppler tracking data, Delta Differential One-way Ranging (Î”DOR) data are also used in the calculation, which shows that they can improve the accuracy of the orbit reconstruction. Total position overlap differences are about 20 m and 30 m for the 100 km*100 km and 100 km*15 km lunar orbit respectively, increased by ~50 % with respect to CE-2. A kinematic statistical method is applied to determine the position of the lander and relative position of the rover with respect to the lander. The location of the lander is computed as: 44.1216Âº N, 19.5124Âº W and -2632.0 m in the lunar Mean Axes coordinate system. The position difference of the lander is better than 50 m compared to the result of the LRO photograph. From 15 to 21 December, the rover walked around the lander, and took photos of each other at the parking point A, B, C, D, E (max distance from the lander is about 25 m). The delta VLBI phase delay data are used to compute the relative position of the rover at the parking points, and the accuracy of the relative position can reach to 1-2 m comparing with the results of visual method.
Yamazaki, Masakazu; Horio, Takuya; Kishimoto, Naoki; Ohno, Koichi
2007-03-15
Although the outer shapes of molecular orbitals (MO's) are of great importance in many phenomena, they have been difficult to be probed by experiments. Here we show that metastable helium (He{sup *}) atoms can sensitively probe the outer properties of molecules and that an electron spectroscopic technique using velocity-selected He{sup *} atoms in combination with classical trajectory simulations leads to a consistent determination of MO functions and the molecular surface. MO functions composed of linear combinations of atomic orbital functions were fitted to the observed collision energy dependences of partial ionization cross sections (CEDPICS). The obtained CEDPICS MO functions were compared with conventionally available Hartree-Fock, Kohn-Sham, and Dyson orbitals.
20 CFR 408.1005 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2011 CFR
2011-04-01
... 20 Employees' Benefits 2 2011-04-01 2011-04-01 false Will we mail you a notice of the initial..., Definitions, and Initial Determinations Â§ 408.1005 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial determination to you at your last known...
20 CFR 408.1005 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Will we mail you a notice of the initial..., Definitions, and Initial Determinations Â§ 408.1005 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial determination to you at your last known...
20 CFR 408.1005 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2014 CFR
2014-04-01
... 20 Employees' Benefits 2 2014-04-01 2014-04-01 false Will we mail you a notice of the initial..., Definitions, and Initial Determinations Â§ 408.1005 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial determination to you at your last known...
20 CFR 408.1005 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2013 CFR
2013-04-01
... 20 Employees' Benefits 2 2013-04-01 2013-04-01 false Will we mail you a notice of the initial..., Definitions, and Initial Determinations Â§ 408.1005 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial determination to you at your last known...
20 CFR 408.1005 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2012 CFR
2012-04-01
... 20 Employees' Benefits 2 2012-04-01 2012-04-01 false Will we mail you a notice of the initial..., Definitions, and Initial Determinations Â§ 408.1005 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial determination to you at your last known...
Single Step to Orbit; a First Step in a Cooperative Space Exploration Initiative
NASA Technical Reports Server (NTRS)
Lusignan, Bruce; Sivalingam, Shivan
1999-01-01
At the end of the Cold War, disarmament planners included a recommendation to ease reduction of the U.S. and Russian aerospace industries by creating cooperative scientific pursuits. The idea was not new, having earlier been suggested by Eisenhower and Khrushchev to reduce the pressure of the "Military Industrial Complex" by undertaking joint space exploration. The Space Exploration Initiative (SEI) proposed at the end of the Cold War by President Bush and Premier Gorbachev was another attempt to ease the disarmament process by giving the bloated war industries something better to do. The engineering talent and the space rockets could be used for peaceful pursuits, notably for going back to the Moon and then on to Mars with human exploration and settlement. At the beginning of this process in 1992 staff of the Stanford Center for International Cooperation in Space attended the International Space University in Canada, met with Russian participants and invited a Russian team to work with us on a joint Stanford-Russian Mars Exploration Study. A CIA student and Airforce and Navy students just happened to join the Stanford course the next year and all students were aware that the leader of the four Russian engineers was well versed in Russian security. But, as long as they did their homework, they were welcome to participate with other students in defining the Mars mission and the three engineers they sent were excellent. At the end of this study we were invited to give a briefing to Dr. Edward Teller at Stanford's Hoover Institution of War and Peace. We were also encouraged to hold a press conference on Capitol Hill to introduce the study to the world. At a pre-conference briefing at the Space Council, we were asked to please remind the press that President Bush had asked for a cooperative exploration proposal not a U.S. alone initiative. The Stanford-Russian study used Russia's Energia launchers, priced at $300 Million each. The mission totaled out to $71.5 Billion, to send a six-person crew to establish a Mars base and return. It was an on going international venture with plans for new crews, base expansion, and extended exploration at every two year opportunity. The $71.5 Billion international approach contrasted with NASA's own 90-day U.S. - alone study that proposed a package topping $500 Billion by some admissions. NASA's approach was also challenged by an internal D.O.E. proposal at much lower cost, described to the Mars Society last year by Lowell Wood and, of course, by Bob Zubrin's "Mars Direct" proposal.
SCD1 Orbit Determination System: Pre-launch preparation, LEOP performance and routine operations
NASA Astrophysics Data System (ADS)
Kuga, Helio Koiti; Rao, Kondapalli Rama
This paper presents a complete overview of the Orbit Determination System (ODS) software developed by the flight dynamics group of the Division of Space Mechanics and Control (DMC) of the Brazilian Institute for Space Research (INPE) for the first Brazilian satellite SCD1. The paper is divided into four parts. The first part explains in brief the SCD1 mission, its ground and space segments and the principal characteristics of its launch system. The second part, i.e. the pre-launch preparation of the software, describes the structure of the ODS adopted for SCD1, and includes a brief history of its development, of its testing with real data of foreign satellites, and of its assessment through the comparison of accuracies obtained. The third part, i.e. the Launch and Early Orbit Phase (LEOP) performance, narrates the experience of the flight dynamics group on the fateful day of the launch: all the odds against the process of orbit determination in terms of lack of enough tracking data, failure of the launch vehicle staff in providing the injection information, last minute modifications of the flight plan, and a few hours of anxiety which preceded the successful follow-up of the mission. The fourth part, i.e. the routine operations part, explains the methodology adopted for using the ODS in day-to-day operations, the accuracy in extended pass-predictions for the Brazilian tracking stations, and the overall performance of the ODS for SCD1. In addition, one also comments about the necessary modifications made during the routine operations along time and possible future improvements to be introduced in the software for the upcoming missions.
20 CFR 404.904 - Notice of the initial determination.
Code of Federal Regulations, 2013 CFR
2013-04-01
... written notice will explain in simple and clear language what we have determined and the reasons for and the effect of our determination. If our determination involves a determination of disability that is... language a statement of the case setting forth the evidence on which our determination is based. The...
20 CFR 416.1404 - Notice of the initial determination.
Code of Federal Regulations, 2013 CFR
2013-04-01
... known address. The written notice will explain in simple and clear language what we have determined and the reasons for and the effect of our determination. If our determination involves a determination of... understandable language a statement of the case setting forth the evidence on which our determination is...
Initial Test Determination of Cosmogenic Nuclides in Magnetite
NASA Astrophysics Data System (ADS)
Matsumura, H.; Caffee, M. W.; Nagao, K.; Nishiizumi, K.
2014-12-01
Long-lived radionuclides, such as 10Be, 26Al, and 36Cl, are produced by cosmic rays in surficial materials on Earth, and used for determinations of cosmic-ray exposure ages and erosion rates. Quartz and limestone are routinely used as the target minerals for these geomorphological studies. Magnetite also contains target elements that produce abundant cosmogenic nuclides when exposed to the cosmic rays. Magnetite has several notable merits that enable the measurement of cosmogenic nuclides: (1) the target elements for production of cosmogenic nuclides in magnetite comprise the dominant mineral form of magnetite, Fe3O4; (2) magnetite can be easily isolated, using a magnet, after rock milling; (3) multiple cosmogenic nuclides are produced by exposure of magnetite to cosmic-ray secondaries; and (4) cosmogenic nuclides produced in the rock containing the magnetite, but not within the magnetite itself, can be separated using nitric acid and sodium hydroxide leaches. As part of this initial study, magnetite was separated from a basaltic sample collected from the Atacama Desert in Chili (2,995 m). Then Be, Al, Cl, Ca, and Mn were separated from ~2 g of the purified magnetite. We measured cosmogenic 10Be, 26Al, and 36Cl concentrations in the magnetite by accelerator mass spectrometry at PRIME Lab, Purdue University. Cosmogenic 3He and 21Ne concentrations of aliquot of the magnetite were measured by mass spectrometry at the University of Tokyo. We also measured the nuclide concentrations from magnetite collected from a mine at Ishpeming, Michigan as a blank. The 10Be and 36Cl concentrations as well as 3He concentration produce concordant cosmic ray exposure ages of ~0.4 Myr for the Atacama basalt. However, observed high 26Al and 21Ne concentrations attribute to those nuclides incorporation from silicate impurity.
Orbit determination accuracy assessment for an asteroid flyby - A Galileo case study
NASA Technical Reports Server (NTRS)
Kechichian, Jean A.; Kenyon, Paul R.; Moultrie, Benjamin
1987-01-01
The Galileo spacecraft may be targeted for a close flyby of an asteroid on its way to an encounter with Jupiter. An orbit determination accuracy analysis was carried out for the case of the asteroid 29 Amphitrite based on the use of radio metric and optical data types. Prior to encounter, the uncertainty in the asteroid's position, based on astrometric observations from earth, amounts to several hundred kilometers. This ephemeris uncertainty constitutes the dominant error in the determination of the spacecraft orbit with respect to Amphitrite. It is shown that the spacecraft-asteroid relative position can be improved by imaging asteroid-star pairs with the Galileo charge-coupled device (CCD) camera, enabling an accurate flyby of the asteroid. The main benefit of optical navigation is to enable the instrument pointing updates necessary for closeup viewing of the asteroid. A discussion of the evolution of the target error ellipse parameters as a function of data coverage and various combinations of radiometric and optical data types is also presented.
Accurate Determination of Comet and Asteroid Orbits Leading to Collision With Earth
NASA Technical Reports Server (NTRS)
Roithmayr, Carlos M.; Kay-Bunnell, Linda; Mazanek, Daniel D.; Kumar, Renjith R.; Seywald, Hans; Hausman, Matthew A.
2005-01-01
Movements of the celestial bodies in our solar system inspired Isaac Newton to work out his profound laws of gravitation and motion; with one or two notable exceptions, all of those objects move as Newton said they would. But normally harmonious orbital motion is accompanied by the risk of collision, which can be cataclysmic. The Earth s moon is thought to have been produced by such an event, and we recently witnessed magnificent bombardments of Jupiter by several pieces of what was once Comet Shoemaker-Levy 9. Other comets or asteroids may have met the Earth with such violence that dinosaurs and other forms of life became extinct; it is this possibility that causes us to ask how the human species might avoid a similar catastrophe, and the answer requires a thorough understanding of orbital motion. The two red square flags with black square centers displayed are internationally recognized as a warning of an impending hurricane. Mariners and coastal residents who know the meaning of this symbol and the signs evident in the sky and ocean can act in advance to try to protect lives and property; someone who is unfamiliar with the warning signs or chooses to ignore them is in much greater jeopardy. Although collisions between Earth and large comets or asteroids occur much less frequently than landfall of a hurricane, it is imperative that we learn to identify the harbingers of such collisions by careful examination of an object s path. An accurate determination of the orbit of a comet or asteroid is necessary in order to know if, when, and where on the Earth s surface a collision will occur. Generally speaking, the longer the warning time, the better the chance of being able to plan and execute action to prevent a collision. The more accurate the determination of an orbit, the less likely such action will be wasted effort or, what is worse, an effort that increases rather than decreases the probability of a collision. Conditions necessary for a collision to occur are discussed, and warning times for long-period comets and near-Earth asteroids are presented.
NASA Astrophysics Data System (ADS)
Bakhshiyan, B. Ts.; Sukhanov, A. A.; Fedyaev, K. S.
2010-10-01
An analysis of the existing astrometric and radar observations of the Apophis asteroid is performed. On the basis of this analysis, characteristics of future measurements of the asteroid orbit and limitation on their conduction are accepted. A proposed launching of a spacecraft to the asteroid in order to obtain high-accuracy measurements of its distance and radial velocity is also considered. Trajectories of the flight to the asteroid in 2012-2022 are studied. Estimates of the accuracy of the Apophis position determination at various sets of both available and planned measurements at various numbers of determined parameters are obtained. The method of estimating accuracy is similar to that used in [1] for the Vega project.
NASA Astrophysics Data System (ADS)
Fang, Haijian; Zhang, Rongzhi; Wang, Jiasong; Wang, Dan; Guo, Hai
2015-10-01
The injected transfer orbit of lunar probe Chang'E 5T1 (CE-5T1) is determined immediately after the probe separates from its launcher. As the first orbit in the lunar flight, the CE-5T1 injected transfer orbit is crucial to the consequence of rocket vehicle launch mission and the probe's subsequent midway orbital manoeuvre. In this paper, we discuss the problem of using rocket GPS measurements to determine the probe velocity increment due to mechanical separation, and subsequently the injected transfer orbit determination of CE-5T1. Motivated by the post-mission analysis of lunar probe Chang'E 3 (CE-3), we give theoretical evidence to explain the physical phenomenon of semi-major axis sudden change at the probe separation instant through the derivation of the Vis-Viva equation. In succession, we focus on the description of the procedure used for the orbit determination performed on separated arcs of rocket GPS measurements through the use of momentum conservation to determine the probe separation velocity. Finally, actual flight data of the CE-3 and CE-5T1 missions are used for the validation.
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2009-11-01
A new method is suggested for finding the preliminary orbit of a small celestial body from its three pairs of angular measurements at three times. The method uses the intermediate orbit that we previously constructed from three position vectors and the corresponding times. This intermediate orbit allows for most of the perturbations in the motion of the body under study. The methodical error of orbit computation by the proposed method is generally three orders smaller than the corresponding error of the traditional approach based on the construction of the unperturbed Keplerian orbit. This fact allows such a reference arc to be selected that the accuracy of the intermediate orbit would always match that of the reference observations that determine this arc. The new method is a highly efficient tool, which allows reliable parameters of the perturbed motion to be obtained already at the stage of computing the preliminary orbit. It is especially efficient if applied to high-accuracy observational data covering short orbital arcs.
NASA Astrophysics Data System (ADS)
Desmars, J.; Camargo, J. I. B.; Braga-Ribas, F.; Vieira-Martins, R.; Assafin, M.; Vachier, F.; Colas, F.; Ortiz, J. L.; Duffard, R.; Morales, N.; Sicardy, B.; Gomes-Júnior, A. R.; Benedetti-Rossi, G.
2015-12-01
Context. The prediction of stellar occultations by trans-Neptunian objects (TNOs) and Centaurs is a difficult challenge that requires accuracy both in the occulted star position and in the object ephemeris. Until now, the most used method of prediction, involving dozens of TNOs/Centaurs, has been to consider a constant offset for the right ascension and for the declination with respect to a reference ephemeris, usually the latest public version. This offset is determined as the difference between the most recent observations of the TNO/Centaur and the reference ephemeris. This method can be successfully applied when the offset remains constant with time, i.e. when the orbit is stable enough. In this case, the prediction even holds for occultations that occur several days after the last observations. Aims: This paper presents an alternative method of prediction, based on a new accurate orbit determination procedure, which uses all the available positions of the TNO from the Minor Planet Center database, as well as sets of new astrometric positions from unpublished observations. Methods: Orbits were determined through a numerical integration procedure called NIMA, in which we developed a specific weighting scheme that considers the individual precision of the observation, the number of observations performed during one night by the same observatory, and the presence of systematic errors in the positions. Results: The NIMA method was applied to 51 selected TNOs and Centaurs. For this purpose, we performed about 2900 new observations in several observatories (European South Observatory, Observatório Pico dos Dias, Pic du Midi, etc.) during the 2007-2014 period. Using NIMA, we succeed in predicting the stellar occultations of 10 TNOs and 3 Centaurs between July 2013 and February 2015. By comparing the NIMA and Jet Propulsion Laboratory (JPL) ephemerides, we highlight the variation in the offset between them with time, by showing that, generally, the constant offset hypothesis is not valid, even for short time scales of a few weeks. Giving examples, we show that the constant offset method cannot accurately predict 6 out of the 13 observed positive occultations that have been successfully predicted by NIMA. The results indicate that NIMA is capable of efficiently refining the orbits of these bodies. Finally, we show that the astrometric positions given by positive occultations can help to refine the orbit of the TNO and, consequently, the future predictions. We also provide unpublished observations of the 51 selected TNOs and their ephemeris in a usable format by the SPICE library. We provide ephemerides of TNO/Centaurs usable with SPICE library and available at http://www.imcce.fr/~desmars/research/tno/The offset observations of the selected TNOs are only available at the CDS via anonymous ftp to http://cdsarc.u-strasbg.fr (ftp://130.79.128.5) or via http://cdsarc.u-strasbg.fr/viz-bin/qcat?J/A+A/584/A96
NASA Astrophysics Data System (ADS)
Son, Ju Young; Jo, Jung Hyun; Choi, Jin
2015-09-01
To protect and manage the Korean space assets including satellites, it is important to have precise positions and orbit information of each space objects. While Korea currently lacks optical observatories dedicated to satellite tracking, the Korea Astronomy and Space Science Institute (KASI) is planning to establish an optical observatory for the active generation of space information. However, due to geopolitical reasons, it is difficult to acquire an adequately sufficient number of optical satellite observatories in Korea. Against this backdrop, this study examined the possible locations for such observatories, and performed simulations to determine the differences in precision of optical orbit estimation results in relation to the relative baseline distance between observatories. To simulate more realistic conditions of optical observation, white noise was introduced to generate observation data, which was then used to investigate the effects of baseline distance between optical observatories and the simulated white noise. We generated the optical observations with white noise to simulate the actual observation, estimated the orbits with several combinations of observation data from the observatories of various baseline differences, and compared the estimated orbits to check the improvement of precision. As a result, the effect of the baseline distance in combined optical GEO satellite observation is obvious but small compared to the observation resolution limit of optical GEO observation.
NASA Technical Reports Server (NTRS)
Hejduk, M. D.; Cowardin, H. M.; Stansbery, Eugene G.
2012-01-01
In performing debris surveys of deep-space orbital regions, the considerable volume of the area to be surveyed and the increased orbital altitude suggest optical telescopes as the most efficient survey instruments; but to proceed this way, methodologies for debris object size estimation using only optical tracking and photometric information are needed. Basic photometry theory indicates that size estimation should be possible if satellite albedo and shape are known. One method for estimating albedo is to try to determine the object's material type photometrically, as one can determine the albedos of common satellite materials in the laboratory. Examination of laboratory filter photometry (using Johnson BVRI filters) on a set of satellite material samples indicates that most material types can be separated at the 1-sigma level via B-R versus R-I color differences with a relatively small amount of required resampling, and objects that remain ambiguous can be resolved by B-R versus B-V color differences and solar radiation pressure differences. To estimate shape, a technique advanced by Hall et al. [1], based on phase-brightness density curves and not requiring any a priori knowledge of attitude, has been modified slightly to try to make it more resistant to the specular characteristics of different materials and to reduce the number of samples necessary to make robust shape determinations. Working from a gallery of idealized debris shapes, the modified technique identifies most shapes within this gallery correctly, also with a relatively small amount of resampling. These results are, of course, based on relatively small laboratory investigations and simulated data, and expanded laboratory experimentation and further investigation with in situ survey measurements will be required in order to assess their actual efficacy under survey conditions; but these techniques show sufficient promise to justify this next level of analysis.
On the Determination of Poisson Statistics for Haystack Radar Observations of Orbital Debris
NASA Technical Reports Server (NTRS)
Stokely, Christopher L.; Benbrook, James R.; Horstman, Matt
2007-01-01
A convenient and powerful method is used to determine if radar detections of orbital debris are observed according to Poisson statistics. This is done by analyzing the time interval between detection events. For Poisson statistics, the probability distribution of the time interval between events is shown to be an exponential distribution. This distribution is a special case of the Erlang distribution that is used in estimating traffic loads on telecommunication networks. Poisson statistics form the basis of many orbital debris models but the statistical basis of these models has not been clearly demonstrated empirically until now. Interestingly, during the fiscal year 2003 observations with the Haystack radar in a fixed staring mode, there are no statistically significant deviations observed from that expected with Poisson statistics, either independent or dependent of altitude or inclination. One would potentially expect some significant clustering of events in time as a result of satellite breakups, but the presence of Poisson statistics indicates that such debris disperse rapidly with respect to Haystack's very narrow radar beam. An exception to Poisson statistics is observed in the months following the intentional breakup of the Fengyun satellite in January 2007.
NASA Astrophysics Data System (ADS)
Guo, Jing; Xu, Xiaolong; Zhao, Qile; Liu, Jingnan
2015-10-01
This contribution summarizes the strategy used by Wuhan University (WHU) to determine precise orbit and clock products for Multi-GNSS Experiment (MGEX) of the International GNSS Service (IGS). In particular, the satellite attitude, phase center corrections, solar radiation pressure model developed and used for BDS satellites are addressed. In addition, this contribution analyzes the orbit and clock quality of the quad-constellation products from MGEX Analysis Centers (ACs) for a common time period of 1 year (2014). With IGS final GPS and GLONASS products as the reference, Multi-GNSS products of WHU (indicated by WUM) show the best agreement among these products from all MGEX ACs in both accuracy and stability. 3D Day Boundary Discontinuities (DBDs) range from 8 to 27 cm for Galileo-IOV satellites among all ACs' products, whereas WUM ones are the largest (about 26.2 cm). Among three types of BDS satellites, MEOs show the smallest DBDs from 10 to 27 cm, whereas the DBDs for all ACs products are at decimeter to meter level for GEOs and one to three decimeter for IGSOs, respectively. As to the satellite laser ranging (SLR) validation for Galileo-IOV satellites, the accuracy evaluated by SLR residuals is at the one decimeter level with the well-known systematic bias of about -5 cm for all ACs. For BDS satellites, the accuracy could reach decimeter level, one decimeter level, and centimeter level for GEOs, IGSOs, and MEOs, respectively. However, there is a noticeable bias in GEO SLR residuals. In addition, systematic errors dependent on orbit angle related to mismodeled solar radiation pressure (SRP) are present for BDS GEOs and IGSOs. The results of Multi-GNSS combined kinematic PPP demonstrate that the best accuracy of position and fastest convergence speed have been achieved using WUM products, particularly in the Up direction. Furthermore, the accuracy of static BDS only PPP degrades when the BDS IGSO and MEO satellites switches to orbit-normal orientation, particularly for COM products, whereas the WUM show the slightest degradation.
NASA Astrophysics Data System (ADS)
Guo, Jing; Xu, Xiaolong; Zhao, Qile; Liu, Jingnan
2016-02-01
This contribution summarizes the strategy used by Wuhan University (WHU) to determine precise orbit and clock products for Multi-GNSS Experiment (MGEX) of the International GNSS Service (IGS). In particular, the satellite attitude, phase center corrections, solar radiation pressure model developed and used for BDS satellites are addressed. In addition, this contribution analyzes the orbit and clock quality of the quad-constellation products from MGEX Analysis Centers (ACs) for a common time period of 1 year (2014). With IGS final GPS and GLONASS products as the reference, Multi-GNSS products of WHU (indicated by WUM) show the best agreement among these products from all MGEX ACs in both accuracy and stability. 3D Day Boundary Discontinuities (DBDs) range from 8 to 27 cm for Galileo-IOV satellites among all ACs' products, whereas WUM ones are the largest (about 26.2 cm). Among three types of BDS satellites, MEOs show the smallest DBDs from 10 to 27 cm, whereas the DBDs for all ACs products are at decimeter to meter level for GEOs and one to three decimeter for IGSOs, respectively. As to the satellite laser ranging (SLR) validation for Galileo-IOV satellites, the accuracy evaluated by SLR residuals is at the one decimeter level with the well-known systematic bias of about -5 cm for all ACs. For BDS satellites, the accuracy could reach decimeter level, one decimeter level, and centimeter level for GEOs, IGSOs, and MEOs, respectively. However, there is a noticeable bias in GEO SLR residuals. In addition, systematic errors dependent on orbit angle related to mismodeled solar radiation pressure (SRP) are present for BDS GEOs and IGSOs. The results of Multi-GNSS combined kinematic PPP demonstrate that the best accuracy of position and fastest convergence speed have been achieved using WUM products, particularly in the Up direction. Furthermore, the accuracy of static BDS only PPP degrades when the BDS IGSO and MEO satellites switches to orbit-normal orientation, particularly for COM products, whereas the WUM show the slightest degradation.
20 CFR 410.620 - Notice of initial determination.
Code of Federal Regulations, 2011 CFR
2011-04-01
... OF 1969, TITLE IV-BLACK LUNG BENEFITS (1969- ) Determinations of Disability, Other Determinations... that a party's entitlement to benefits has ended because of such party's death (see § 410.610(c))....
GPS-Based Precision Orbit Determination for a New Era of Altimeter Satellites: Jason-1 and ICESat
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.; Rowlands, David D.; Lemoine, Frank G.; Zelensky, Nikita P.; Williams, Teresa A.
2003-01-01
Accurate positioning of the satellite center of mass is necessary in meeting an altimeter mission's science goals. The fundamental science observation is an altimetric derived topographic height. Errors in positioning the satellite's center of mass directly impact this fundamental observation. Therefore, orbit error is a critical Component in the error budget of altimeter satellites. With the launch of the Jason-1 radar altimeter (Dec. 2001) and the ICESat laser altimeter (Jan. 2003) a new era of satellite altimetry has begun. Both missions pose several challenges for precision orbit determination (POD). The Jason-1 radial orbit accuracy goal is 1 cm, while ICESat (600 km) at a much lower altitude than Jason-1 (1300 km), has a radial orbit accuracy requirement of less than 5 cm. Fortunately, Jason-1 and ICESat POD can rely on near continuous tracking data from the dual frequency codeless BlackJack GPS receiver and Satellite Laser Ranging. Analysis of current GPS-based solution performance indicates the l-cm radial orbit accuracy goal is being met for Jason-1, while radial orbit accuracy for ICESat is well below the 54x1 mission requirement. A brief overview of the GPS precision orbit determination methodology and results for both Jason-1 and ICESat are presented.
NASA Astrophysics Data System (ADS)
Lala, P.
1981-04-01
Papers are presented on the use of point mass models of the geopotential for orbit predictions, on earth ocean tides from long-term analysis of satellite orbits, on the motion of an artificial satellite under the terrestrial radiation pressure, and on the generation of satellite position (and velocity) by a mixed analytical-numerical procedure. Attention is also given to the possibilities of determining the influence of earth body tides on the motion of artificial satellites and to a general time element for orbit integration in Cartesian coordinates.
42 CFR 405.927 - Initial determinations subject to the reopenings process.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 42 Public Health 2 2010-10-01 2010-10-01 false Initial determinations subject to the reopenings... HEALTH AND HUMAN SERVICES MEDICARE PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED...) Initial Determinations § 405.927 Initial determinations subject to the reopenings process. Minor errors...
42 CFR 405.927 - Initial determinations subject to the reopenings process.
Code of Federal Regulations, 2014 CFR
2014-10-01
... 42 Public Health 2 2014-10-01 2014-10-01 false Initial determinations subject to the reopenings... HEALTH AND HUMAN SERVICES MEDICARE PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED...) Initial Determinations § 405.927 Initial determinations subject to the reopenings process. Minor errors...
NASA Astrophysics Data System (ADS)
Bogdanova, Iu. P.; Semenov, B. I.
1989-03-01
A method for determining the orbit of a space object from angular measurements is proposed which is based on the use of the theory of multipoint boundary value problems for the set of ordinary nonlinear differential equations describing the motion of the object in the terrestrial spheroid field. The proposed approach makes it possible to synthesize algorithms for finding the orbit of a space object according to a minimum number of measurements.
Numerical comparison of discrete Kalman filter algorithms - Orbit determination case study
NASA Technical Reports Server (NTRS)
Bierman, G. J.; Thornton, C. L.
1976-01-01
Numerical characteristics of various Kalman filter algorithms are illustrated with a realistic orbit determination study. The case study of this paper highlights the numerical deficiencies of the conventional and stabilized Kalman algorithms. Computational errors associated with these algorithms are found to be so large as to obscure important mismodeling effects and thus cause misleading estimates of filter accuracy. The positive result of this study is that the U-D covariance factorization algorithm has excellent numerical properties and is computationally efficient, having CPU costs that differ negligibly from the conventional Kalman costs. Accuracies of the U-D filter using single precision arithmetic consistently match the double precision reference results. Numerical stability of the U-D filter is further demonstrated by its insensitivity to variations in the a priori statistics.
A numerical comparison of discrete Kalman filtering algorithms: An orbit determination case study
NASA Technical Reports Server (NTRS)
Thornton, C. L.; Bierman, G. J.
1976-01-01
The numerical stability and accuracy of various Kalman filter algorithms are thoroughly studied. Numerical results and conclusions are based on a realistic planetary approach orbit determination study. The case study results of this report highlight the numerical instability of the conventional and stabilized Kalman algorithms. Numerical errors associated with these algorithms can be so large as to obscure important mismodeling effects and thus give misleading estimates of filter accuracy. The positive result of this study is that the Bierman-Thornton U-D covariance factorization algorithm is computationally efficient, with CPU costs that differ negligibly from the conventional Kalman costs. In addition, accuracy of the U-D filter using single-precision arithmetic consistently matches the double-precision reference results. Numerical stability of the U-D filter is further demonstrated by its insensitivity of variations in the a priori statistics.
Orbit determination by solving for gravity parameters with multiple arc data
NASA Technical Reports Server (NTRS)
Wu, Jiun-Tsong
1992-01-01
The orbit of a satellite that repeats in the earth fixed coordinates is determined by combining GPS tracking data from multiple arcs. The satellite dynamics are modeled with the epoch state and a set of parameters, called the bin parameters, that account for the effect of the local gravitational field on the satellite current state. The epoch state is specific to each arc, and the bin parameters are common to all repeat arcs. The estimation algorithm is based on the Square Root Information Filter. It involves partitioning of the measurement matrix and use of the Householder transformation to combine multiple arc data and solve for the epoch states and the bin parameters. The bin parameters can then be converted into the earth's gravitational field with a modest amount of computation.
Enhanced orbit determination filter: Inclusion of ground system errors as filter parameters
NASA Technical Reports Server (NTRS)
Masters, W. C.; Scheeres, D. J.; Thurman, S. W.
1994-01-01
The theoretical aspects of an orbit determination filter that incorporates ground-system error sources as model parameters for use in interplanetary navigation are presented in this article. This filter, which is derived from sequential filtering theory, allows a systematic treatment of errors in calibrations of transmission media, station locations, and earth orientation models associated with ground-based radio metric data, in addition to the modeling of the spacecraft dynamics. The discussion includes a mathematical description of the filter and an analytical comparison of its characteristics with more traditional filtering techniques used in this application. The analysis in this article shows that this filter has the potential to generate navigation products of substantially greater accuracy than more traditional filtering procedures.
Determination of On-Orbit Cabin Air Loss from the International Space Station (ISS)
NASA Technical Reports Server (NTRS)
Williams, David E.; Leonard, Daniel J.; Smith, Patrick J.
2004-01-01
The International Space Station (ISS) loses cabin atmosphere mass at some rate. Due to oxygen partial pressures fluctuations from metabolic usage, the total pressure is not a good data source for tracking total pressure loss. Using the nitrogen partial pressure is a good data source to determine the total on-orbit cabin atmosphere loss from the ISS, due to no nitrogen addition or losses. There are several important reasons to know the daily average cabin air loss of the ISS including logistics planning for nitrogen and oxygen. The total average daily cabin atmosphere loss was estimated from January 14 to April 9 of 2003. The total average daily cabin atmosphere loss includes structural leakages, Vozdukh losses, Carbon Dioxide Removal Assembly (CDRA) losses, and other component losses. The total average daily cabin atmosphere loss does not include mass lost during Extra-Vehicular Activities (EVAs), Progress dockings, Space Shuttle dockings, calibrations, or other specific one-time events.
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.; Zelensky, N. P.; Rowlands, D. D.; Lemoine, F. G.; Chinn, D. S.; Williams, T. A.
2002-01-01
Jason-1, launched on December 7,2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. T P has demonstrated that the time variation of ocean topography can be determined with an accuracy of a few centimeters, thanks to the availability of highly accurate orbits based primarily on SLR+DORIS tracking. The Jason-1 mission is intended to continue measurement of the ocean surface with the same, if not better accuracy. Fortunately, Jason- 1 POD can rely on four independent tracking data types available including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. Orbit solutions computed using individual and various combinations of GPS, SLR, DORIS and altimeter crossover data types have been determined from over 100 days of Jason-1 tracking data, The performance of the orbit solutions and tracking data has been evaluated. Orbit solution evaluation and comparison has provided insight into possible areas of refinement. Several aspects of the POD process are examined to obtain orbit improvements including measurement modeling, force modeling and solution strategy. The results of these analyses will be presented.
Determinants of breastfeeding initiation among mothers in Kuwait
2010-01-01
Background Exclusive breastfeeding is recommended as the optimal way to feed infants for the first six months of life. While overall breastfeeding rates are high, exclusive breastfeeding is relatively uncommon among Middle Eastern women. The objective of this study was to identify the incidence of breastfeeding amongst women in the six governorates of Kuwait and the factors associated with the initiation of breastfeeding. Methods A sample of 373 women (aged 17-47 years), recruited shortly after delivery from four hospitals in Kuwait, completed a structured, interviewer-administered questionnaire. Multivariate logistic regression analysis was used to identify those factors independently associated with the initiation of breastfeeding. Results In total, 92.5% of mothers initiated breastfeeding and at discharge from hospital the majority of mothers were partially breastfeeding (55%), with only 30% of mothers fully breastfeeding. Prelacteal feeding was the norm (81.8%) and less than 1 in 5 infants (18.2%) received colostrum as their first feed. Only 10.5% of infants had been exclusively breastfed since birth, the remainder of the breastfed infants having received either prelacteal or supplementary infant formula feeds at some time during their hospital stay. Of the mothers who attempted to breastfeed, the majority of women (55.4%) delayed their first attempt to breastfeed until 24 hours or more after delivery. Breastfeeding at discharge from hospital was positively associated with paternal support for breastfeeding and negatively associated with delivery by caesarean section and with the infant having spent time in the Special Care Nursery. Conclusions The reasons for the high use of prelacteal and supplementary formula feeding warrant investigation. Hospital policies and staff training are needed to promote the early initiation of breastfeeding and to discourage the unnecessary use of infant formula in hospital, in order to support the establishment of exclusive breastfeeding by mothers in Kuwait. PMID:20667112
Code of Federal Regulations, 2010 CFR
2010-07-01
... MULTIEMPLOYER PLANS ALLOCATING UNFUNDED VESTED BENEFITS TO WITHDRAWING EMPLOYERS Allocation Methods for Merged... two or more plans that use the presumptive allocation method of section 4211(b) of ERISA merge, the merged plan may adjust the amortization of initial liabilities under § 4211.32(b) to amortize...
Code of Federal Regulations, 2011 CFR
2011-07-01
... MULTIEMPLOYER PLANS ALLOCATING UNFUNDED VESTED BENEFITS TO WITHDRAWING EMPLOYERS Allocation Methods for Merged... two or more plans that use the presumptive allocation method of section 4211(b) of ERISA merge, the merged plan may adjust the amortization of initial liabilities under § 4211.32(b) to amortize...
Code of Federal Regulations, 2013 CFR
2013-07-01
... MULTIEMPLOYER PLANS ALLOCATING UNFUNDED VESTED BENEFITS TO WITHDRAWING EMPLOYERS Allocation Methods for Merged... two or more plans that use the presumptive allocation method of section 4211(b) of ERISA merge, the merged plan may adjust the amortization of initial liabilities under § 4211.32(b) to amortize...
Code of Federal Regulations, 2014 CFR
2014-07-01
... MULTIEMPLOYER PLANS ALLOCATING UNFUNDED VESTED BENEFITS TO WITHDRAWING EMPLOYERS Allocation Methods for Merged... two or more plans that use the presumptive allocation method of section 4211(b) of ERISA merge, the merged plan may adjust the amortization of initial liabilities under § 4211.32(b) to amortize...
Code of Federal Regulations, 2012 CFR
2012-07-01
... MULTIEMPLOYER PLANS ALLOCATING UNFUNDED VESTED BENEFITS TO WITHDRAWING EMPLOYERS Allocation Methods for Merged... two or more plans that use the presumptive allocation method of section 4211(b) of ERISA merge, the merged plan may adjust the amortization of initial liabilities under § 4211.32(b) to amortize...
NASA Technical Reports Server (NTRS)
Lemoine, Frank G.; Zelensky, Nikita P.; Chinn, Douglas S.; Beckley, Brian D.; Lillibridge, John L.
2006-01-01
The US Navy's GEOSAT Follow-On spacecraft (GFO) primary mission objective is to map the oceans using a radar altimeter. Satellite laser ranging data, especially in combination with altimeter crossover data, offer the only means of determining high-quality precise orbits. Two tuned gravity models, PGS7727 and PGS7777b, were created at NASA GSFC for GFO that reduce the predicted radial orbit through degree 70 to 13.7 and 10.0 mm. A macromodel was developed to model the nonconservative forces and the SLR spacecraft measurement offset was adjusted to remove a mean bias. Using these improved models, satellite-ranging data, altimeter crossover data, and Doppler data are used to compute both daily medium precision orbits with a latency of less than 24 hours. Final precise orbits are also computed using these tracking data and exported with a latency of three to four weeks to NOAA for use on the GFO Geophysical Data Records (GDR s). The estimated orbit precision of the daily orbits is between 10 and 20 cm, whereas the precise orbits have a precision of 5 cm.
Determination of the Venus flyby orbits of the Soviet Vega probes using VLBI techniques
NASA Technical Reports Server (NTRS)
Ellis, J.; Mcelrath, Timothy P.
1988-01-01
In December 1984, the Soviet Union launched two identical Vega spacecraft with the dual objective of exploring Venus and continuing to rendezvous with the comet Halley. The two Vega spacecraft encountered Venus in mid-June 1985 and successfully deployed entry probes and wind-measuring balloons into the Venus atmosphere. An objective of the Venus Balloon experiment was to measure the Venus winds using differential VLBI from the balloon and the flyby bus. NASA's Deep Space 64 meter subnet was part of a world wide network organized to collect data from the Vega probes and balloons. A critical element of this experiment was an accurate determination of the Venus relative flyby orbits of the Vega spacecraft during the 46 hour balloon lifetime. Venus flyby solutions were independently determined by the Soviets using two-way range and Doppler from Soviet stations and by JPL using one-way Doppler and VLBI data collected from the DSN. The Vega flyby solutions determined by the Soviets using a sparse two-way tracking strategy with JPL solutions using the DSN VLBI data to complement the Soviet data and with solutions using only one-way data collected by the DSN were compared.
NASA Astrophysics Data System (ADS)
Zhou, ShanShi; Hu, XiaoGong; Wu, Bin; Liu, Li; Qu, WeiJing; Guo, Rui; He, Feng; Cao, YueLing; Wu, XiaoLi; Zhu, LingFeng; Shi, Xin; Tan, HongLi
2011-06-01
Aiming at regional services, the space segment of COMPASS (Phase I) satellite navigation system is a constellation of Geostationary Earth Orbit (GEO), Inclined Geostationary Earth Orbit (IGSO) and Medium Earth Orbit (MEO) satellites. Precise orbit determination (POD) for the satellites is limited by the geographic distribution of regional tracking stations. Independent time synchronization (TS) system is developed to supplement the regional tracking network, and satellite clock errors and orbit data may be obtained by simultaneously processing both tracking data and TS data. Consequently, inconsistency between tracking system and TS system caused by remaining instrumental errors not calibrated may decrease navigation accuracy. On the other hand, POD for the mixed constellation of GEO/IGSO/MEO with the regional tracking network leads to parameter estimations that are highly correlated. Notorious example of correlation is found between GEO's orbital elements and its clock errors. We estimate orbital elements and clock errors for a 3GEO+2IGSO constellation in this study using a multi-satellite precise orbit determination (MPOD) strategy, with which clock error elimination algorithm is applied to separate orbital and clock estimates to improve numerical efficiency. Satellite Laser Ranging (SLR) data are used to evaluate User Ranging Error (URE), which is the orbital error projected on a receiver's line-of-sight direction. Two-way radio-wave time transfer measurements are used to evaluate clock errors. Experimenting with data from the regional tracking network, we conclude that the fitting of code data is better than 1 m in terms of Root-Mean-Square (RMS), and fitting of carrier phase is better than 1 cm. For orbital evaluation, difference between computed receiver-satellite ranging based on estimated orbits and SLR measurements is better than 1 m (RMS). For clock estimates evaluation, 2-hour linear-fitting shows that the satellite clock rates are about 1.E-10 s/s, while receiver clock rates are about 1×10-13-1×10-12 s/s. For the 72-hour POD experiment, the average differences between POD satellite clock rates estimates and clock measurements based on TS system are about 1×10-13 s/s, and for receiver clock rates, the differences are about 1×10-15 s/s.
24 CFR 599.301 - Initial determination of threshold requirements.
Code of Federal Regulations, 2011 CFR
2011-04-01
... requirements. Before rating and ranking an application, HUD will review it to determine if the application... requirements of § 599.107. (b) Failure to meet threshold requirements—(1) No rating or ranking. An...
24 CFR 599.301 - Initial determination of threshold requirements.
Code of Federal Regulations, 2014 CFR
2014-04-01
... requirements. Before rating and ranking an application, HUD will review it to determine if the application... requirements of § 599.107. (b) Failure to meet threshold requirements—(1) No rating or ranking. An...
NASA Technical Reports Server (NTRS)
Throckmorton, D. A.
1982-01-01
Temperatures measured at the aerodynamic surface of the Orbiter's thermal protection system (TPS), and calorimeter measurements, are used to determine heating rates to the TPS surface during atmospheric entry. On the Orbiter leeside, where convective heating rates are low, it is possible that a significant portion of the total energy input may result from solar radiation, and for the wing, cross radiation from the hot (relatively) Orbiter fuselage. In order to account for the potential impact of these sources, values of solar- and cross-radiation heat transfer are computed, based upon vehicle trajectory and attitude information and measured surface temperatures. Leeside heat-transfer data from the STS-2 mission are presented, and the significance of solar radiation and fuselage-to-wing cross-radiation contributions to total energy input to Orbiter leeside surfaces is assessed.
Lewis, Karen M.; Fujii, Yuka
2014-08-20
We survey the methods proposed in the literature for detecting moons of extrasolar planets in terms of their ability to distinguish between prograde and retrograde moon orbits, an important tracer of the moon formation channel. We find that most moon detection methods, in particular, sensitive methods for detecting moons of transiting planets, cannot observationally distinguishing prograde and retrograde moon orbits. The prograde and retrograde cases can only be distinguished where the dynamical evolution of the orbit due to, e.g., three body effects is detectable, where one of the two cases is dynamically unstable, or where new observational facilities, which can implement a technique capable of differentiating the two cases, come online. In particular, directly imaged planets are promising targets because repeated spectral and photometric measurements, which are required to determine moon orbit direction, could also be conducted with the primary interest of characterizing the planet itself.
Determinants of initiation and progression of idiopathic pulmonary fibrosis.
Kottmann, Robert Matthew; Hogan, Christopher M; Phipps, Richard P; Sime, Patricia J
2009-09-01
IPF is a devastating disease with few therapeutic options. The precise aetiology of IPF remains elusive. However, our understanding of the pathologic processes involved in the initiation and progression of this disease is improving. Data on the mechanisms underlying IPF have been generated from epidemiologic investigations as well as cellular and molecular studies of human tissues. Although no perfect animal model of human IPF exists, pre-clinical animal studies have helped define pathways which are likely important in human disease. Epithelial injury, fibroblast activation and repetitive cycles of injury and abnormal repair are almost certainly key events. Factors which have been associated with initiation and/or progression of IPF include viral infections, abnormal cytokine, chemokine and growth factor production, oxidant stress, autoimmunity, inhalational of toxicants and gastro-oesophageal reflux disease. Furthermore, recent evidence identifies a role for a variety of genetic and epigenetic abnormalities ranging from mutations in surfactant protein C to abnormalities in telomere length and telomerase activity. The challenge remains to identify additional inciting agents and key dysregulated pathways that lead to disease progression so that we can develop targeted therapies to treat or prevent this serious disease. PMID:19740254
Determinants of initiation and progression of idiopathic pulmonary fibrosis
Kottmann, Robert Matthew; Hogan, Christopher M.; Phipps, Richard P.; Sime, Patricia J.
2013-01-01
IPF is a devastating disease with few therapeutic options. The precise aetiology of IPF remains elusive. However, our understanding of the pathologic processes involved in the initiation and progression of this disease is improving. Data on the mechanisms underlying IPF have been generated from epidemiologic investigations as well as cellular and molecular studies of human tissues. Although no perfect animal model of human IPF exists, pre-clinical animal studies have helped define pathways which are likely important in human disease. Epithelial injury, fibroblast activation and repetitive cycles of injury and abnormal repair are almost certainly key events. Factors which have been associated with initiation and/or progression of IPF include viral infections, abnormal cytokine, chemokine and growth factor production, oxidant stress, autoimmunity, inhalational of toxicants and gastro-oesophageal reflux disease. Furthermore, recent evidence identifies a role for a variety of genetic and epigenetic abnormalities ranging from mutations in surfactant protein C to abnormalities in telomere length and telomerase activity. The challenge remains to identify additional inciting agents and key dysregulated pathways that lead to disease progression so that we can develop targeted therapies to treat or prevent this serious disease. PMID:19740254
NASA Astrophysics Data System (ADS)
Hackel, Stefan; Montenbruck, Oliver; Steigenberger, -Peter; Eineder, Michael; Gisinger, Christoph
Remote sensing satellites support a broad range of scientific and commercial applications. The two radar imaging satellites TerraSAR-X and TanDEM-X provide spaceborne Synthetic Aperture Radar (SAR) and interferometric SAR data with a very high accuracy. The increasing demand for precise radar products relies on sophisticated validation methods, which require precise and accurate orbit products. Basically, the precise reconstruction of the satellite’s trajectory is based on the Global Positioning System (GPS) measurements from a geodetic-grade dual-frequency receiver onboard the spacecraft. The Reduced Dynamic Orbit Determination (RDOD) approach utilizes models for the gravitational and non-gravitational forces. Following a proper analysis of the orbit quality, systematics in the orbit products have been identified, which reflect deficits in the non-gravitational force models. A detailed satellite macro model is introduced to describe the geometry and the optical surface properties of the satellite. Two major non-gravitational forces are the direct and the indirect Solar Radiation Pressure (SRP). Due to the dusk-dawn orbit configuration of TerraSAR-X, the satellite is almost constantly illuminated by the Sun. Therefore, the direct SRP has an effect on the lateral stability of the determined orbit. The indirect effect of the solar radiation principally contributes to the Earth Radiation Pressure (ERP). The resulting force depends on the sunlight, which is reflected by the illuminated Earth surface in the visible, and the emission of the Earth body in the infrared spectra. Both components of ERP require Earth models to describe the optical properties of the Earth surface. Therefore, the influence of different Earth models on the orbit quality is assessed within the presentation. The presentation highlights the influence of non-gravitational force and satellite macro models on the orbit quality of TerraSAR-X.
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2012 CFR
2012-10-01
... following: (1) A determination with respect to entitlement to hospital insurance or supplementary medical insurance under Medicare. (2) A disallowance of an individual's application for entitlement to hospital or... of an application for hospital or supplementary medical insurance, or a denial of a request...
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
... following: (1) A determination with respect to entitlement to hospital insurance or supplementary medical insurance under Medicare. (2) A disallowance of an individual's application for entitlement to hospital or... of an application for hospital or supplementary medical insurance, or a denial of a request...
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2013 CFR
2013-10-01
... following: (1) A determination with respect to entitlement to hospital insurance or supplementary medical insurance under Medicare. (2) A disallowance of an individual's application for entitlement to hospital or... of an application for hospital or supplementary medical insurance, or a denial of a request...
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2011 CFR
2011-10-01
... following: (1) A determination with respect to entitlement to hospital insurance or supplementary medical insurance under Medicare. (2) A disallowance of an individual's application for entitlement to hospital or... of an application for hospital or supplementary medical insurance, or a denial of a request...
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2014 CFR
2014-10-01
... following: (1) A determination with respect to entitlement to hospital insurance or supplementary medical insurance under Medicare. (2) A disallowance of an individual's application for entitlement to hospital or... of an application for hospital or supplementary medical insurance, or a denial of a request...
12 CFR 792.59 - Appeal of initial determination.
Code of Federal Regulations, 2012 CFR
2012-01-01
... jurisdiction of the Office of Personnel Management, appeals will be made pursuant to that agency's regulations... appropriate. (e) If access is denied because of an exemption, the individual will be notified of the right to... information to which the amendment or statement of disagreement relates whose identity can be determined by...
12 CFR 792.59 - Appeal of initial determination.
Code of Federal Regulations, 2011 CFR
2011-01-01
... jurisdiction of the Office of Personnel Management, appeals will be made pursuant to that agency's regulations... appropriate. (e) If access is denied because of an exemption, the individual will be notified of the right to... information to which the amendment or statement of disagreement relates whose identity can be determined by...
Orbital angular momentum in electron diffraction and its use to determine chiral crystal symmetries
NASA Astrophysics Data System (ADS)
Juchtmans, Roeland; Verbeeck, Jo
2015-10-01
In this work we present an alternative way to look at electron diffraction in a transmission electron microscope. Instead of writing the scattering amplitude in Fourier space as a set of plane waves, we use the cylindrical Fourier transform to describe the scattering amplitude in a basis of orbital angular momentum (OAM) eigenstates. We show how working in this framework can be very convenient when investigating, e.g., rotation and screw-axis symmetries. For the latter we find selection rules on the OAM coefficients that unambiguously reveal the handedness of the screw axis. Detecting the OAM coefficients of the scattering amplitude thus offers the possibility to detect the handedness of crystals without the need for dynamical simulations, the thickness of the sample, nor the exact crystal structure. We propose an experimental setup to measure the OAM components where an image of the crystal is taken after inserting a spiral phase plate in the diffraction plane and perform multislice simulations on ? quartz to demonstrate how the method indeed reveals the chirality. The experimental feasibility of the technique is discussed together with its main advantages with respect to chirality determination of screw axes. The method shows how the use of a spiral phase plate can be extended from a simple phase imaging technique to a tool to measure the local OAM decomposition of an electron wave, widening the field of interest well beyond chiral space group determination.
GPS interferometric attitude and heading determination - Initial flight test results
NASA Technical Reports Server (NTRS)
Van Graas, Frank; Braasch, Michael
1992-01-01
Attitude and heading determination using GPS interferometry is a well-understood concept. However, efforts have been concentrated mainly in the development of robust algorithms and applications for low-dynamic, rigid platforms (e.g., shipboard). This paper presents results of what is believed to be the first real-time flight test of a GPS attitude and heading determination system. Signals from four antennas are processed by a 24-channel GPS receiver. Data from the receiver are sent to a microcomputer for storage and further computations. Attitude and heading data are sent to a second computer for display on a software-generated artificial horizon. Demonstration of this technique proves its candidacy for augmentation of aircraft state estimation for flight control and navigation, as well as for numerous other applications.
GPS interferometric attitude and heading determination: Initial flight test results
NASA Technical Reports Server (NTRS)
Vangraas, Frank; Braasch, Michael
1991-01-01
Attitude and heading determination using GPS interferometry is a well-understood concept. However, efforts have been concentrated mainly in the development of robust algorithms and applications for low dynamic, rigid platforms (e.g., shipboard). This paper presents results of what is believed by the authors to be the first realtime flight test of a GPS attitude and heading determination system. The system is installed in Ohio University's Douglas DC-3 research aircraft. Signals from four antennas are processed by an Ashtech 3DF 24-channel GPS receiver. Data from the receiver are sent to a microcomputer for storage and further computations. Attitude and heading data are sent to a second computer for display on a software generated artificial horizon. Demonstration of this technique proves its candidacy for augmentation of aircraft state estimation for flight control and navigation as well as for numerous other applications.
NASA Technical Reports Server (NTRS)
Smith, R. L.; Huang, C.
1986-01-01
A recent mathematical technique for solving systems of equations is applied in a very general way to the orbit determination problem. The study of this technique, the homotopy continuation method, was motivated by the possible need to perform early orbit determination with the Tracking and Data Relay Satellite System (TDRSS), using range and Doppler tracking alone. Basically, a set of six tracking observations is continuously transformed from a set with known solution to the given set of observations with unknown solutions, and the corresponding orbit state vector is followed from the a priori estimate to the solutions. A numerical algorithm for following the state vector is developed and described in detail. Numerical examples using both real and simulated TDRSS tracking are given. A prototype early orbit determination algorithm for possible use in TDRSS orbit operations was extensively tested, and the results are described. Preliminary studies of two extensions of the method are discussed: generalization to a least-squares formulation and generalization to an exhaustive global method.
DPTRAJ/ODP - DOUBLE PRECISION TRAJECTORY ANALYSIS AND ORBIT DETERMINATION PROGRAM
NASA Technical Reports Server (NTRS)
Breckheimer, P. J.
1994-01-01
The Double Precision Trajectory Analysis Program, DPTRAJ, and the Orbit Determination Program, ODP, have been developed and improved over the years to provide the NASA Jet Propulsion Laboratory with a highly reliable and accurate navigation capability for their deep space missions such as VOYAGER. DPTRAJ and ODP are each collections of programs which work together to provide the desired computational results. DPTRAJ, ODP, and their supporting utility programs are capable of handling the massive amounts of data and performing the various numerical calculations required for solving the navigation problems associated with planetary fly-by and lander missions. They were used extensively in support of NASA's VOYAGER project. DPTRAJ produces a spacecraft ephemeris by numerical integration of the equations of motion, which can be formulated using a full set of acceleration models. For each particular trajectory case the extent of the modeling employed and the precision of the integration process are controlled by user input specifications. The equation of motion used includes four types of terms. An acceleration term accounts for the basic conic motion of the spacecraft with respect to the central body. Terms that measure the attraction of the perturbing bodies on the spacecraft and terms that indirectly affect the motion as perturbations on the central body may be included. Terms are also provided to account for other gravitational and non-gravitational effects on the motion. ODP's function is the processing of the observational data in order to compute precise estimates of the spacecraft, or lander, position coordinate histories. This function is executed by processing the observation data and auxiliary calibration information. ODP also computes a spacecraft state vector, or a lander position vector, along with parameters which define the acceleration. The heart of the ODP process is a data fitting subprocess in which validated, edited, and corrected observational data is transformed into a state vector estimate. The derived state vector estimate may then be used to generate an estimated trajectory. This trajectory contains the final product of the orbit determination process, which is the time evolution of the estimated spacecraft, or lander, position coordinates. DPTRAJ-ODP is written in FORTRAN V, SFTRAN, PL/I and ASSEMBLER for use on DEC VAX series computers running VMS, and has a central memory requirement of 3.4Mb. This program is available on a 1600 BPI 9-track magnetic tape in VAX BACKUP format. DPTRAJ and ODP were originally developed on a UNIVAC 1100 series computer. The VAX/VMS version was developed in 1987.
Orbit Determination of the Mars Global Surveyor Spacecraft Using Laser Altimetry
NASA Technical Reports Server (NTRS)
Smith, David E.; Zuber, M. T.; Lemoine, F. G.; Rowlands, D. D.
2001-01-01
Many of the scientific investigations of the Mars Global Surveyor (MGS) mission require high precision orbital information and some are limited entirely by its quality. These include the laser altimeter (MOLA) the Mars gravity field and atmospheric occultation investigations by radio science, and the planetary dynamics and celestial mechanics investigations. The precision of the orbits can usually be assessed by comparing overlapping orbits for a given period; but these results tend to reflect the repeatability rather than the accuracy. The re-constructed orbits from the doppler and range tracking data on MGS are (to date) at the few meter level radially, and a few hundreds of meters horizontally, using the best gravity models, presently available. With the laser altimeter on MGS we have a mechanism to measure the quality and to actually make significant improvements in the orbital accuracy by incorporating the altimetry data as a tracking datatype. By adding the altimeter measurements at orbital cross-over locations we have been able to reduce die radial error to 1 meter of less on average and have reduced the along track and out of plane error by almost 2 orders of magnitude down to a few meters. It is apparent that the altimeter observation provides a geometric strength to the orbit that it is not possible to obtain from the present doppler and the range data alone. We discuss the results obtained for the first year of the MGS mapping orbit. This work is supported by the NASA Mars Program.
NASA Astrophysics Data System (ADS)
Montenbruck, Oliver
1991-02-01
It is shown that Taylor series integration allows problems of celestial mechanics for interplanetary orbits to be solved with relatively simple orbit models. The method is adapted to the computation of satellite orbits. A new implementation concept is included, which allows a programmation of the equations of motion. It offers simultaneously extension possibilities, which are necessary for the computation of special terms of the power function. The linking of Sun and Moon ephemeris for the treatment of gavitational disturbances of the satellite orbit is used as an example. This concept is represented with the principles of Taylor series integration, and compared with variants of the process. The power spectrum of the method is examined for disturbed and undisturbed Kepler orbits.
NASA Astrophysics Data System (ADS)
Chapront, J.; Chapront-Touzé, M.; Francou, G.
2002-05-01
An analysis of Lunar Laser Ranging (LLR) observations from January 1972 until April 2001 has been performed, and a new solution for the lunar orbital motion and librations has been constructed that has been named S2001. With respect to prior solutions, improvements in the statistical treatment of the data, new nutation and libration models and the addition of the positions of the observing stations to the list of fitted parameters have been introduced. Globally, for recent observations, our rms (root mean square error) is within 2 to 3 centimeters in the lunar distance. Special attention has been paid to the determination of the correction to the IAU76 luni-solar constant of precession, and the value of the secular acceleration of the Moon's longitude due to the tidal forces. The main results are: - correction to the constant of precession: Delta p = -0.302 +/- 0.003 ''/cy, - tidal acceleration of the lunar longitude: Gamma = -25.858 +/- 0.003 ''/cy2. The positions and velocities of the stations have also been determined. The results are consistent with the ITRF2000 determinations from SLR observations. The lunar theory ELP is referred to a dynamical system and introduces the inertial mean ecliptic of J2000.0. The positioning of the reference system of the theory with respect to ICRS is performed (and also with respect to some useful JPL numerical integrations). Finally the orientation of the celestial axes with respect to the ICRS reference system has been derived as well as the offsets of the Celestial Ephemeris Pole.
NASA Technical Reports Server (NTRS)
Radomski, M. S.; Doll, C. E.
1991-01-01
This investigation concerns the effects on Ocean Topography Experiment (TOPEX) spacecraft operational orbit determination of ionospheric refraction error affecting tracking measurements from the Tracking and Data Relay Satellite System (TDRSS). Although tracking error from this source is mitigated by the high frequencies (K-band) used for the space-to-ground links and by the high altitudes for the space-to-space links, these effects are of concern for the relatively high-altitude (1334 kilometers) TOPEX mission. This concern is due to the accuracy required for operational orbit-determination by the Goddard Space Flight Center (GSFC) and to the expectation that solar activity will still be relatively high at TOPEX launch in mid-1992. The ionospheric refraction error on S-band space-to-space links was calculated by a prototype observation-correction algorithm using the Bent model of ionosphere electron densities implemented in the context of the Goddard Trajectory Determination System (GTDS). Orbit determination error was evaluated by comparing parallel TOPEX orbit solutions, applying and omitting the correction, using the same simulated TDRSS tracking observations. The tracking scenarios simulated those planned for the observation phase of the TOPEX mission, with a preponderance of one-way return-link Doppler measurements. The results of the analysis showed most TOPEX operational accuracy requirements to be little affected by space-to-space ionospheric error. The determination of along-track velocity changes after ground-track adjustment maneuvers, however, is significantly affected when compared with the stringent 0.1-millimeter-per-second accuracy requirements, assuming uncoupled premaneuver and postmaneuver orbit determination. Space-to-space ionospheric refraction on the 24-hour postmaneuver arc alone causes 0.2 millimeter-per-second errors in along-track delta-v determination using uncoupled solutions. Coupling the premaneuver and postmaneuver solutions, however, appears likely to reduce this figure substantially. Plans and recommendations for response to these findings are presented.
NASA Astrophysics Data System (ADS)
W?odarczyk, K.; W?odarczyk, I.
2014-07-01
A detailed analysis of the passage through the atmosphere of a very bright meteor that exploded in the air near Chelyabinsk, Russia on February 15, 2013 is presented. A number of videos and photographs were examined thoroughly to determine the meteor trajectory beginning from the recorded atmospheric entry height of about 62.5 km until its disappearance at about 9.8 km. The calculated velocity changes as a function time revealed an unusual behavior: during the first 10 seconds the meteor velocity increased from 16.6 km/s up to about 20.6 km/s in the main air burst at the altitude of 26.5 km. Afterwards it decreased rapidly. The light curves derived from videos enabled the total radiant energy and mass loss variations to be calculated. The heliocentric orbit of the meteoroid and possible parent bodies were computed. We proposed an additional 'close approaches' method to the existing method of checking meteoroid/bolide parent bodies based on different D-criteria.
NASA Technical Reports Server (NTRS)
Ulvestad, J. S.; Thurman, S. W.
1992-01-01
An error covariance analysis methodology is used to investigate different weighting schemes for two-way (coherent) Doppler data in the presence of transmission-media and observing-platform calibration errors. The analysis focuses on orbit-determination performance in the interplanetary cruise phase of deep-space missions. Analytical models for the Doppler observable and for transmission-media and observing-platform calibration errors are presented, drawn primarily from previous work. Previously published analytical models were improved upon by the following: (1) considering the effects of errors in the calibration of radio signal propagation through the troposphere and ionosphere as well as station-location errors; (2) modelling the spacecraft state transition matrix using a more accurate piecewise-linear approximation to represent the evolution of the spacecraft trajectory; and (3) incorporating Doppler data weighting functions that are functions of elevation angle, which reduce the sensitivity of the estimated spacecraft trajectory to troposphere and ionosphere calibration errors. The analysis is motivated by the need to develop suitable weighting functions for two-way Doppler data acquired at 8.4 GHz (X-band) and 32 GHz (Ka-band). This weighting is likely to be different from that in the weighting functions currently in use; the current functions were constructed originally for use with 2.3 GHz (S-band) Doppler data, which are affected much more strongly by the ionosphere than are the higher frequency data.
NASA Technical Reports Server (NTRS)
Lemoine, Frank G.; Rowlands, David D.; Luthcke, Scott B.; Zelensky, Nikita P.; Chinn, Douglas S.; Pavlis, Despina E.; Marr, Gregory
2001-01-01
The US Navy's GEOSAT Follow-On Spacecraft was launched on February 10, 1998 with the primary objective of the mission to map the oceans using a radar altimeter. Following an extensive set of calibration campaigns in 1999 and 2000, the US Navy formally accepted delivery of the satellite on November 29, 2000. Satellite laser ranging (SLR) and Doppler (Tranet-style) beacons track the spacecraft. Although limited amounts of GPS data were obtained, the primary mode of tracking remains satellite laser ranging. The GFO altimeter measurements are highly precise, with orbit error the largest component in the error budget. We have tuned the non-conservative force model for GFO and the gravity model using SLR, Doppler and altimeter crossover data sampled over one year. Gravity covariance projections to 70x70 show the radial orbit error on GEOSAT was reduced from 2.6 cm in EGM96 to 1.3 cm with the addition of SLR, GFO/GFO and TOPEX/GFO crossover data. Evaluation of the gravity fields using SLR and crossover data support the covariance projections and also show a dramatic reduction in geographically-correlated error for the tuned fields. In this paper, we report on progress in orbit determination for GFO using GFO/GFO and TOPEX/GFO altimeter crossovers. We will discuss improvements in satellite force modeling and orbit determination strategy, which allows reduction in GFO radial orbit error from 10-15 cm to better than 5 cm.
Determining the type of orbits in the central regions of barred galaxies
NASA Astrophysics Data System (ADS)
Zotos, Euaggelos E.; Caranicolas, Nicolaos D.
2016-02-01
We use a simple dynamical model which consists of a harmonic oscillator and a spherical component, in order to investigate the regular or chaotic character of orbits in a barred galaxy with a central spherically symmetric nucleus. Our aim is to explore how the basic parameters of the galactic system influence the nature of orbits, by computing in each case the percentage of chaotic orbits, as well as the percentages of different types of regular orbits. We also give emphasis to the types of regular orbits that support either the formation of nuclear rings or the barred structure of the galaxy. We provide evidence that the traditional x1 orbital family does not always dominate in barred galaxy models since we found several other types of resonant orbits which can also support the barred structure. We also found that sparse enough nuclei, fast rotating bars and high energy models can support the galactic bars. On the other hand, weak bars, dense central nuclei, slowly rotating bars and low energy models favor the formation of nuclear rings. We also compare our results with previous related work.
NASA Astrophysics Data System (ADS)
Klimyk, Anatoliy; Patera, Jiri
2007-02-01
In the paper, properties of antisymmetric orbit functions are reviewed and further developed. Antisymmetric orbit functions on the Euclidean space En are antisymmetrized exponential functions. Antisymmetrization is fulfilled by a Weyl group, corresponding to a Coxeter-Dynkin diagram. Properties of such functions are described. These functions are closely related to irreducible characters of a compact semisimple Lie group G of rank n. Up to a sign, values of antisymmetric orbit functions are repeated on copies of the fundamental domain F of the affine Weyl group (determined by the initial Weyl group) in the entire Euclidean space En. Antisymmetric orbit functions are solutions of the corresponding Laplace equation in En, vanishing on the boundary of the fundamental domain F. Antisymmetric orbit functions determine a so-called antisymmetrized Fourier transform which is clo! sely related to expansions of central functions in characters of irreducible representations of the group G. They also determine a transform on a finite set of points of F (the discrete antisymmetric orbit function transform). Symmetric and antisymmetric multivariate exponential, sine and cosine discrete transforms are given.
The Cooling Rate of an Active Aa Lava Flow Determined Using an Orbital Imaging Spectrometer
NASA Astrophysics Data System (ADS)
Wright, Robert; Garbeil, Harold
2010-05-01
The surface temperature of an active lava flow is an important physical property to measure. Through its influence on lava crystallinity, cooling exerts a fundamental control on lava rheology. Remotely sensed thermal radiance data acquired by multispectral sensors such as Landsat Thematic Mapper and the Terra Advanced Spaceborne Thermal Emission and Reflection Radiometer, are of insufficient spectral and radiometric fidelity to allow for realistic determination of lava surface temperatures from Earth orbit. This paper presents results obtained from the analysis of active lava flows using hyperspectral data acquired by NASA's Earth Observing-1 Hyperion imaging spectrometer. The contiguous nature of the measured radiance spectrum in the 0.4-2.5 micron region means that, although sensor saturation most certainly occurs, unsaturated radiance data are always available from even the hottest, and most radiant, active lava flow surfaces. The increased number of wavebands available allows for the assumption of more complex flow surface temperature distributions in the radiance-to-temperature inversion processes. The technique is illustrated by using a hyperspectral image of the active lava lake at Erta Ale volcano, Ethiopia, a well characterized calibration target. We then go on to demonstrate how this approach can be used to constrain the surface cooling rate of an active lava flow at Mount Etna, Sicily, using three images acquired during a four day period in September 2004. The cooling rate of the active channel as determined from space falls within the limits commonly assumed in numerical lava flow models. The results provide insights into the temperature-radiance mixture modeling problem that will aid in the analysis of data acquired by future hyperspectral remote sensing missions, such as NASA's proposed HyspIRI mission.
NASA Astrophysics Data System (ADS)
Todd, Paul; Pierson, Duane L.; Allen, Britt; Silverstein, JoAnn
The formation of biofilms by water microorganisms such as Pseudomonas aeruginosa in spacecraft water systems has been a matter of concern for long-duration space flight. Crewed spacecraft plumbing includes internal surfaces made of 316L stainless steel. Experiments were therefore undertaken to compare the ability of P. aeruginosa to grow in suspension, attach to stainless steel and to grow on stainless steel in low gravity on the space shuttle. Four categories of cultures were studied during two space shuttle flights (STS-69 and STS-77). Cultures on the ground were held in static horizontal or vertical cylindrical containers or were tumbled on a clinostat and activated under conditions identical to those for the flown cultures. The containers used on the ground and in flight were BioServe Space Technologiesâ€™ Fluid Processing Apparatus (FPA), an open-ended test tube with rubber septa that allows robotic addition of bacteria to culture media to initiate experiments and the addition of fixative to conclude experiments. Planktonic growth was monitored by spectrophotometry, and biofilms were characterized quantitatively by epifluorescence and scanning electron microscopy. In these experiments it was found that: (1) Planktonic growth in flown cultures was more extensive than in static cultures, as seen repeatedly in the history of space microbiology, and closely resembled the growth of tumbled cultures. (2) Conversely, the attachment of cells in flown cultures was as much as 8 times that in tumbled cultures but not significantly different from that in static horizontal and vertical cultures, consistent with the notion that flowing fluid reduces microbial attachment. (3) The final surface coverage in 8 days was the same for flown and static cultures but less by a factor of 15 in tumbled cultures, where coverage declined during the preceding 4 days. It is concluded that cell attachment to 316L stainless steel in the low gravity of orbital space flight is similar to that found in stagnant cultures at 1 x g. Research was supported by NASA contract NAGW-1197 to the University of Colorado.
Orbit determination results and trajectory reconstruction for the Cassini/Huygens Mission
NASA Technical Reports Server (NTRS)
Bordi, John J.; Antreasian, Pete; Jones, Jerry; Meek, Cameron; Ionasescu, Rodica; Roundhill, Ian; Roth, Duane
2005-01-01
During Cassini's third orbit around Saturn, the Huygens Probe was successfully released on a trajectory that resulted in the probe entering Titan's atmosphere on January 14, 2005, making it both the most distant spacecraft landing and the first spacecraft to successfully land on the moon of another planet. This paper documents the reconstruction of both the orbiter and probe trajectoriespanning the Titan-B and Titan-C encounters.
Composition of the Moon as Determined from Orbit by Gamma-Ray Spectroscopy
NASA Technical Reports Server (NTRS)
Metzger, A. E.
1994-01-01
A spacecraft placed in a planetary orbit of suitably high inclination will pass over all or most of the planet's surface in a matter of several weeks to months. The quite prodigious scientific potential of planetary orbiters lies in coupling this comprehensive coverage with observing systems capable of gathering data on properties that include elemental and mineralogic composition, exogenic and endogenic surface alterations, thermal balance, gravity, topography, stratigraphy, albedo and magnetism.
Nonlinear Observability for Relative Orbit Determination with Angles-Only Measurements
NASA Astrophysics Data System (ADS)
Kaufman, Evan; Lovell, T. Alan; Lee, Taeyoung
2016-02-01
This paper presents nonlinear observability criteria for the relative orbital dynamics represented by the solutions of the two-body problem. It is assumed that a chief is on a circular orbit with a prescribed orbital radius, and it measures lines-of-sight toward a deputy only. A differential geometric method, based on the Lie derivatives, is used to derive sufficient conditions for observability of the orbital properties of the deputy. It is shown that under certain geometric conditions on the relative configuration between the chief and the deputy, the nonlinear relative motion is observable from angles-only measurements. The second part of this paper presents a quantitative measure of observability for the relative orbits, and it is formulated by generalizing the observability Gramian of linear dynamic systems. An extended Kalman filter is also developed to numerically illustrate the observability of nonlinear relative orbits with angles-only measurements and to show correspondence between the proposed observability measure and filtered solution accuracy.
NASA Astrophysics Data System (ADS)
Modenini, D.; Tortora, P.
2014-07-01
The present work describes our investigation of the navigation anomaly of the Pioneer 10 and 11 probes which became known as the Pioneer Anomaly. It appeared as a linear drift in the Doppler data received by the spacecraft, which has been ascribed to an approximately constant Sunward acceleration of about 8.5×10-13 km/s2. Since then, the existence of the anomaly has been confirmed independently by several groups and a large effort was devoted to find its origin. Recently, different analyses were published where the authors claimed the acceleration due to anisotropic thermal emission to be the most likely cause of the unexplained acceleration. Here we report the methodology and the results of an independent study carried out in the last years, aimed at supporting the thermal origin of the anomaly. This work consists of two main parts: thermal modeling of the spacecraft throughout its trajectory, and orbit determination analysis. Based on existing documentation and published telemetry data, we built a thermal finite element model of the spacecraft, whose complexity has been constrained to a degree allowing for sensitivity analysis, leading to the computation of its formal uncertainty. The trajectory analysis and orbit determination were carried out using NASA/JPL's Orbit Determination Program, and our results show that orbital solutions are achieved that do not require the addition of any "unknown" acceleration other than that of thermal origin.
NASA Astrophysics Data System (ADS)
de Sanctis, M. L.; Politis, M.-F.; Vuilleumier, R.; Stia, C. R.; Fojón, O. A.
2015-08-01
We theoretically study the single ionization of liquid water by energetic electrons through one active-electron first-order model. We analyze the angular ejected electron spectra corresponding to the most external orbitals 1B1, 2A1, 1B2 and 1A1 of a single water molecule. We work to create a realistic description of those orbitals corresponding to single molecules in the liquid phase. This goal is achieved by means of a Wannier orbital formalism. Multiple differential cross sections are computed and compared with previous calculations for both liquid and gas phases. In addition, our present results are integrated over all orientations and compared with experimental ones for randomly oriented vapour water molecules, as no experiments currently exist for the liquid phase. Moreover, we estimate the influence of the passive electrons on the reaction by means of a model potential.
NASA Astrophysics Data System (ADS)
Li, XiaoJie; Zhou, JianHua; Hu, XiaoGong; Liu, Li; Guo, Rui; Zhou, ShanShi
2015-08-01
Geostationary (GEO) satellites form an indispensable component of the constellation of Beidou navigation system (BDS). The ephemerides, or predicted orbits of these GEO satellites(GEOs), are broadcast to positioning, navigation, and timing users. User equivalent ranging error (UERE) based on broadcast message is better than 1.5 m (root formal errors: RMS) for GEO satellites. However, monitoring of UERE indicates that the orbital prediction precision is significantly degraded when the Sun is close to the Earth's equatorial plane (or near spring or autumn Equinox). Error source analysis shows that the complicated solar radiation pressure on satellite buses and the simple box-wing model maybe the major contributor to the deterioration of orbital precision. With the aid of BDS' two-way frequency and time transfer between the GEOs and Beidou time (BDT, that is maintained at the master control station), we propose a new orbit determination strategy, namely three-step approach of the multi-satellite precise orbit determination (MPOD). Pseudo-range (carrier phase) data are transformed to geometric range (biased geometric range) data without clock offsets; and reasonable empirical acceleration parameters are estimated along with orbital elements to account for the error in solar radiation pressure modeling. Experiments with Beidou data show that using the proposed approach, the GEOs' UERE when near the autumn Equinox of 2012 can be improved to 1.3 m from 2.5 m (RMS), and the probability of user equivalent range error (UERE)<2.0 m can be improved from 50% to above 85%.
NASA Technical Reports Server (NTRS)
Luthcke, Scott; Rowlands, David; Lemoine, Frank; Zelensky, Nikita; Beckley, Brian; Klosko, Steve; Chinn, Doug
2006-01-01
Although satellite altimetry has been around for thirty years, the last fifteen beginning with the launch of TOPEX/Poseidon (TP) have yielded an abundance of significant results including: monitoring of ENS0 events, detection of internal tides, determination of accurate global tides, unambiguous delineation of Rossby waves and their propagation characteristics, accurate determination of geostrophic currents, and a multi-decadal time series of mean sea level trend and dynamic ocean topography variability. While the high level of accuracy being achieved is a result of both instrument maturity and the quality of models and correction algorithms applied to the data, improving the quality of the Climate Data Records produced from altimetry is highly dependent on concurrent progress being made in fields such as orbit determination. The precision orbits form the reference frame from which the radar altimeter observations are made. Therefore, the accuracy of the altimetric mapping is limited to a great extent by the accuracy to which a satellite orbit can be computed. The TP mission represents the first time that the radial component of an altimeter orbit was routinely computed with an accuracy of 2-cm. Recently it has been demonstrated that it is possible to compute the radial component of Jason orbits with an accuracy of better than 1-cm. Additionally, still further improvements in TP orbits are being achieved with new techniques and algorithms largely developed from combined Jason and TP data analysis. While these recent POD achievements are impressive, the new accuracies are now revealing subtle systematic orbit error that manifest as both intra and inter annual ocean topography errors. Additionally the construction of inter-decadal time series of climate data records requires the removal of systematic differences across multiple missions. Current and future efforts must focus on the understanding and reduction of these errors in order to generate a complete and consistent time series of improved orbits across multiple missions and decades required for the most stringent climate-related research. This presentation discusses the POD progress and achievements made over nearly three decades, and presents the future challenges, goals and their impact on altimetric derived ocean sciences.
Cooling rate of some active lavas determined using an orbital imaging spectrometer
NASA Astrophysics Data System (ADS)
Wright, Robert; Garbeil, Harold; Davies, Ashley G.
2010-06-01
The surface temperature of an active lava flow is an important physical property to measure. Through its influence on lava crystallinity, cooling exerts a fundamental control on lava rheology. Remotely sensed thermal radiance data acquired by multispectral sensors such as Landsat Thematic Mapper and the Terra Advanced Spaceborne Thermal Emission and Reflection Radiometer are of insufficient spectral and radiometric fidelity to allow for realistic determination of lava surface temperatures from Earth orbit. This paper presents results obtained from the analysis of active lava flows using hyperspectral data acquired by NASA's Earth Observing-1 Hyperion imaging spectrometer. The contiguous nature of the measured radiance spectrum in the 0.4-2.5 Î¼m region means that, although sensor saturation most certainly occurs, unsaturated radiance data are always available from even the hottest, and most radiant, active lava flow surfaces. The increased number of wave bands available allows for the assumption of more complex flow surface temperature distributions in the radiance-to-temperature inversion processes. The technique is illustrated by using a hyperspectral image of the active lava lake at Erta Ale volcano, Ethiopia, a well-characterized calibration target, a time series of three Hyperion images of an active lava flow acquired during a 4 day period at Mount Etna, Sicily, as well as a lava flow erupted at Nyamuragira, Democratic Republic of Congo. The results provide insights into the temperature-radiance mixture modeling problem that will aid in the analysis of data acquired by future hyperspectral remote sensing missions, such as NASA's proposed HyspIRI mission.
The cooling rate of an aa lava flow determined using an orbital imaging spectrometer
NASA Astrophysics Data System (ADS)
Wright, R.; Garbeil, H.
2009-12-01
The surface temperature of an active lava flow is an important property to measure. Through its influence on lava crystallinity, cooling exerts a fundamental control on lava rheology. Remotely sensed thermal radiance data acquired by multispectral sensors such as Landsat Thematic Mapper and the Terra Advanced Spaceborne Thermal Emission and Reflection Radiometer, are of insufficient spectral and radiometric fidelity to allow for realistic determination of lava surface temperatures from Earth orbit. This presentation describes results obtained from the analysis of active lava flows using hyperspectral data acquired by NASAâ€™s Earth Observing-1 Hyperion imaging spectrometer. The contiguous nature of the measured radiance spectrum in the 0.4-2.5 micron region means that, although sensor saturation most certainly occurs, unsaturated radiance data are always available from even the hottest, and most radiant, active lava flow surfaces. The increased number of wavebands available allows for the assumption of more complex flow surface temperature distributions in the radiance-to-temperature inversion processes. The application of such data to the analysis of a time-series of three Hyperion images of an active lava flow, acquired during a four day period at Mount Etna, Sicily, is demonstrated. The results provide insights into the temperature-radiance mixture modeling problem that will aid in the analysis of data acquired by future hyperspectral remote sensing missions, such as NASAâ€™s proposed HyspIRI mission. By also proving radiance data on the opposite limb of the planckian emittance curve (i.e. the MIR and TIR), HyspIRI will allow us to improve upon these antecedent results.
NASA Astrophysics Data System (ADS)
Tien, J. Y.; Young, L.; Meehan, T.; Franklin, G.; Hurst, K. J.; Esterhuizen, S.; Trig Gnss Receiver Team
2010-12-01
Several planned NASA and NOAA missions require an advanced science-quality GNSS receiver as mission-critical payloads to meet science objectives (e. g. CLARREO, COSMIC-2, ICESat II and DESDynI). The science measurement and mission navigation needs require that GNSS receivers track signals from GPS, GALILEO, and other new GNSS systems. JPL is developing the next generation multi-antenna GNSS receiver called the TriG (Tri-GNSS) Receiver for spaceborne scientific measurements that will enable both the continued access of NASA missions to precision orbit determination for remote sensing missions and the application of GNSS signals for the technically demanding radio occultation observations. The TriG receiver will track both the legacy L1CA, L2 Codeless, and the new L2C and L5 signals from GPS as well as new GNSS signals from Galileo and GLONASS. The ability to track multiple GNSS satellite signals would allow full capability to operate during the transition to GPS-III and past the 2020 retirement of the legacy signals and also significantly improve the quality and quantity of the radio occultation measurements. In addition, the TriG receiver features several innovations including digital beam steering, wideband open loop tracking, and “Blue Shift” signal processing algorithm that would enable the necessary precision in the atmosphere and to increase the SNR from the lower regions of the atmosphere in order to dramatically improve the percentage of profiles reaching into the lowest regions of the atmosphere. This presentation will describe the TriG architecture and features how those may be beneficial for the next-generation of global network instruments.
20 CFR 418.1310 - When may you request that we make a new initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false When may you request that we make a new... Administrative Review Process § 418.1310 When may you request that we make a new initial determination? (a) You may request that we make a new initial determination in the following circumstances: (1) You provide...
20 CFR 418.1310 - When may you request that we make a new initial determination?
Code of Federal Regulations, 2011 CFR
2011-04-01
... 20 Employees' Benefits 2 2011-04-01 2011-04-01 false When may you request that we make a new... Administrative Review Process § 418.1310 When may you request that we make a new initial determination? (a) You may request that we make a new initial determination in the following circumstances: (1) You provide...
40 CFR 142.11 - Initial determination of primary enforcement responsibility.
Code of Federal Regulations, 2010 CFR
2010-07-01
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40 CFR 142.11 - Initial determination of primary enforcement responsibility.
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2014-07-01
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40 CFR 142.11 - Initial determination of primary enforcement responsibility.
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2011-07-01
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7 CFR 1951.706 - Initial determination that unauthorized assistance was received.
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2010-01-01
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37 CFR 102.10 - Appeals from initial determinations or untimely delays.
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2014-07-01
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37 CFR 102.10 - Appeals from initial determinations or untimely delays.
Code of Federal Regulations, 2010 CFR
2010-07-01
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11 CFR 1.9 - Appeal of initial adverse agency determination on amendment or correction.
Code of Federal Regulations, 2013 CFR
2013-01-01
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11 CFR 1.9 - Appeal of initial adverse agency determination on amendment or correction.
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2011-01-01
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11 CFR 1.9 - Appeal of initial adverse agency determination on amendment or correction.
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2014-01-01
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11 CFR 1.9 - Appeal of initial adverse agency determination on amendment or correction.
Code of Federal Regulations, 2012 CFR
2012-01-01
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20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2014 CFR
2014-04-01
... 20 Employees' Benefits 2 2014-04-01 2014-04-01 false Will we mail you a notice of the initial... Medicare Part D Subsidies Determinations and the Administrative Review Process Â§ 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the...
20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2012 CFR
2012-04-01
... 20 Employees' Benefits 2 2012-04-01 2012-04-01 false Will we mail you a notice of the initial... Medicare Part D Subsidies Determinations and the Administrative Review Process Â§ 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the...
20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Will we mail you a notice of the initial... Medicare Part D Subsidies Determinations and the Administrative Review Process Â§ 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the...
20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2013 CFR
2013-04-01
... 20 Employees' Benefits 2 2013-04-01 2013-04-01 false Will we mail you a notice of the initial... Medicare Part D Subsidies Determinations and the Administrative Review Process Â§ 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the...
20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2011 CFR
2011-04-01
... 20 Employees' Benefits 2 2011-04-01 2011-04-01 false Will we mail you a notice of the initial... Medicare Part D Subsidies Determinations and the Administrative Review Process Â§ 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the...
Lifetimes of lunar satellite orbits
NASA Technical Reports Server (NTRS)
Meyer, Kurt W.; Buglia, James J.; Desai, Prasun N.
1994-01-01
The Space Exploration Initiative has generated a renewed interest in lunar mission planning. The lunar missions currently under study, unlike the Apollo missions, involve long stay times. Several lunar gravity models have been formulated, but mission planners do not have enough confidence in the proposed models to conduct detailed studies of missions with long stay times. In this report, a particular lunar gravitational model, the Ferrari 5 x 5 model, was chosen to determine the lifetimes for 100-km and 300-km perilune altitude, near-circular parking orbits. The need to analyze orbital lifetimes for a large number of initial orbital parameters was the motivation for the formulation of a simplified gravitational model from the original model. Using this model, orbital lifetimes were found to be heavily dependent on the initial conditions of the nearly circular orbits, particularly the initial inclination and argument of perilune. This selected model yielded lifetime predictions of less than 40 days for some orbits, and other orbits had lifetimes exceeding a year. Although inconsistencies and limitations are inherent in all existing lunar gravity models, primarily because of a lack of information about the far side of the moon, the methods presented in this analysis are suitable for incorporating the moon's nonspherical gravitational effects on the preliminary design level for future lunar mission planning.
Precise orbit determination for GEOSAT and GEOSAT Follow-On in the GRACE era
NASA Astrophysics Data System (ADS)
Zelensky, N. P.; Lemoine, F.; Beckley, B. D.; Chinn, D. S.; Rowlands, D. D.; Lillibridge, J. L.; Scharroo, R.; Smith, W. H.
2009-12-01
The U.S. Navy GEOSAT mission provided the first long-term altimetric record for studies of ocean circulation, marine gravity/bathymetry and continental ice, from early 1985 through 1989. The GEOSAT Follow-On spacecraft (GFO), launched in 1998, began continuous radar altimeter coverage of the oceans in 2000 and was terminated in late 2008. By providing high quality altimeter data, GEOSAT delivered the first and only altimetric measurements over the 1980’s. GFO supplements Jason, TOPEX/Poseidon (T/P), and Envisat, providing a different synoptic sampling of the oceans with its 17-day ground track repeat cycle. Altimeter crossover analysis suggests GFO and GEOSAT are capable of Poseidon class altimetry, both showing crossover residuals averaging below 7.5 cm, with 5-cm orbit error the largest contributor to the altimeter error budget. This study evaluates improvements to the recently released GEOSAT GGM02C GDR orbits and current GFO GDR orbits. The POD model improvements include application of standards consistent with the latest generation of GSFC reprocessed T/P, Jason-1, and Jason-2 MEaSURES orbits, such as the LPOD2005 set of SLR station coordinates, the EIGEN-GL04S1 gravity model, and improved modeling of the time-varying geopotential using GRACE data and the GOT4.7 ocean tide model. In addition the Doppler station coordinates have been re-estimated. In this presentation we summarize the quality of the orbits and the status of our research effort.
Determination of the area and mass distribution of orbital debris fragments
NASA Technical Reports Server (NTRS)
Badhwar, G. D.; Anz-Meador, P. D.
1988-01-01
An important factor in modeling the orbital debris environment is the loss rate of debris due to atmospheric drag and lunisolar perturbations. An accurate knowledge of the area-to-mass ratio of debris fragments is required to calculate the effects of atmospheric drag. It is shown here that the orbital elements as a function of time can be used to invert any propagation algorithm to yield the area-to-mass ratio of an orbiting object. From these calculations and the observed radar cross-section of the object, the mass can be calculated to an accuracy of about 30 percent. It is shown that the mass is related to the effective cross-section area by a power-law relation, but for a given area the mass distribution is very broad. An expression is given for the cumulative mass distribution.
Computer program PRIOR used for orbit determination at the Ondrejov Observatory
NASA Astrophysics Data System (ADS)
Lala, P.
The PRIOR (Program for Improvement of Orbits) computer program, intended mainly for computation of the orbits of Intercosmos satellites, is described. Attention is given to input data, the coordinate system, the computation of perturbations, and the method of solution. The tests that the program has undergone are described, and tables are included that show the effect of changing the weights of observations and the effect of changing the number of Tesseral harmonics. The results of the calculations carried out for the Geos B and Intercosmos 17 satellites are given.
The Use of Laser Altimetry in the Orbit and Attitude Determination of Mars Global Surveyor
NASA Technical Reports Server (NTRS)
Rowlands, D. D.; Pavlis, D. E.; Lemoine, F. G.; Neumann, G. A.; Luthcke, S. B.
1999-01-01
Altimetry from the Mars Observer Laser Altimeter (MOLA) which is carried on board Mars Global Surveyor (MGS) has been analyzed for the period of the MOS mission known as Science Phasing Orbit 1 (SPO-1). We have used these altimeter ranges to improve orbit and attitude knowledge for MGS. This has been accomplished by writing crossover constraint equations that have been derived from short passes of MOLA data. These constraint equations differ from traditional Crossover constraints and exploit the small foot print associated with laser altimetry.
NASA Technical Reports Server (NTRS)
Kramer, Leonard
2014-01-01
A plasma diagnostic package is deployed on the International Space Station (ISS). The system - a Floating Potential Measurement Unit (FPMU) - is used by NASA to monitor the electrical floating potential of the vehicle to assure astronaut safety during extravehicular activity. However, data from the unit also reflects the ionosphere state and seems to represent an unutilized scientific resource in the form of an archive of scientific plasma state data. The unit comprises a Floating Potential probe and two Langmuir probes. There is also an unused but active plasma impedance probe. The data, at one second cadence, are collected, typically for a two week period surrounding extravehicular activity events. Data is also collected any time a visiting vehicle docks with ISS and also when any large solar events occur. The telemetry system is unusual because the package is mounted on a television camera stanchion and its data is impressed on a video signal that is transmitted to the ground and streamed by internet to two off center laboratory locations. The data quality has in the past been challenged by weaknesses in the integrated ground station and distribution systems. These issues, since mid-2010, have been largely resolved and the ground stations have been upgraded. Downstream data reduction has been developed using physics based modeling of the electron and ion collecting character in the plasma. Recursive algorithms determine plasma density and temperature from the raw Langmuir probe current voltage sweeps and this is made available in real time for situational awareness. The purpose of this paper is to describe and record the algorithm for data reduction and to show that the Floating probe and Langmuir probes are capable of providing long term plasma state measurement in the ionosphere. Geophysical features such as the Appleton anomaly and high latitude modulation at the edge of the Auroral zones are regularly observed in the nearly circular, 51 deg inclined, 400 km altitude ISS orbit. Evidence of waves in the ion collection current data is seen in geographic zones known to exhibit the spread-F phenomenon. An anomaly in the current collection characteristic of the cylindrical probe appears also too be organized by the geomagnetic field.
NASA Technical Reports Server (NTRS)
Morinelli, Patrick J.; Ward, Douglas T.; Blizzard, Michael R.; Mendelsohn, Chad R.
2008-01-01
This paper provides an overview of the lessons learned from the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center s (GSFC) Flight Dynamics Facility s (FDF) support of the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft emergency in February 2007, and the Tracking and Data Relay Satellite-3 (TDRS-3) spacecraft emergency in March 2006. A successful and timely recovery from both of these spacecraft emergencies depended on accurate knowledge of the orbit. Unfortunately, the combination of each spacecraft emergency with very little tracking data contributed to difficulties in estimating and predicting the orbit and delayed recovery efforts in both cases. In both the THEMIS and TDRS-3 spacecraft emergencies, numerous factors contributed to problems with obtaining nominal tracking data measurements. This paper details the various causative factors and challenges. This paper further enumerates lessons learned from FDF s recovery efforts involving the THEMIS and TDRS-3 spacecraft emergencies and scant tracking data, as well as recommendations for improvements and corrective actions. In addition, this paper describes the broad range of resources and complex navigation methods employed within the FDF for supporting critical navigation activities during all mission phases, including launch, early orbit, and on-orbit operations.