A new method of initial orbit determination
B. G. Morton; L. G. Taff
1986-01-01
A new method of initial orbit determination for angles-only data is presented. The technique is applicable with only three data sets but the formalism can incorporate arbitrarily large amounts of data. The algorithm rests on the fact that the orbital plane is usually very well determined, as a consequence of the central nature of the gravitational force, and a theorem
A new method of initial orbit determination
B. G. Morton; L. G. Taff
1986-01-01
We present a new method of initial orbit determination for angles-only data. The technique is applicable with only three data sets but the formalism can incorporate arbitrarily large amounts of data. The algorithm rests on the fact that the orbital plane is usually very well determined, as a consequence of the central nature of the gravitational force, and a theorem
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.
A Comprehensive Comparison Between Angles-Only Initial Orbit Determination Techniques
Schaeperkoetter, Andrew Vernon
2012-02-14
. . . . . . . . . . . . . . . . . . . . . . 51 2. Polar Orbit . . . . . . . . . . . . . . . . . . . . . . . . 52 viii CHAPTER Page 3. Sun-Synchronous Orbit . . . . . . . . . . . . . . . . . 54 4. Molniya Orbit . . . . . . . . . . . . . . . . . . . . . . 55 5. GEO Orbit... : : : : : : : : : : : : : : : : : 69 22 E ects of Initial Guesses on Orbital Estimate for Perigee on Molniya Orbit Using Double R Method : : : : : : : : : : : : : : : : : 70 23 E ects of Initial Guesses on Orbital Estimate for GEO Orbit Using Double R Method...
Performance Evaluation of Orbit Determination System during Initial Phase of INSAT-3 Mission
NASA Astrophysics Data System (ADS)
Subramanian, B.; Vighnesam, N. V.
INSAT-3C is the second in the third generation of ISRO's INSAT series of satellites that was launched by ARIANE-SPACE on 23 January 2002 at 23 h 46 m 57 s (lift off time in U.T). The ARIANE-4 Flight Nr.147 took off from Kourou in French Guyana and injected the 2750-kg communications satellite in a geostationary transfer orbit of (571 X 35935) km with an inclination of 4.007 deg at 00 h 07 m 48 s U.T on 24 January 2002 (1252 s after lift off). The satellite was successfully guided into its intended geostationary position of 74 deg E longitude by 09 February 2002 after a series of four firings of its Liquid Apogee Motor (LAM) and four station acquisition (STAQ) maneuvers. Six distinct phases of the mission were categorized based on the orbit characteristics of the INSAT- 3C mission, namely, the pre-launch phase, the launch phase, transfer orbit phase, intermediate orbit phase, drift orbit phase and synchronous orbit phase. The orbit with a perigee height of 571 km at injection of the satellite, was gradually raised to higher orbits with perigee height increasing to 9346 km after Apogee Motor Firing #1 (AMF #1), 18335 km after AMF #2, 32448 km after AMF #3 and 35493 km after AMF #4. The North and South solar panels and the reflectors were deployed at this stage of the mission and the attitude of the satellite with respect to the three axes was stabilized. The Orbit Determination System (ODS) that was used in the initial phase of the mission played a crucial role in realizing the objectives of the mission. This system which consisted of Tracking Data Pre-Processing (TDPP) software, Ephemeris Generation (EPHGEN) software and the Orbit Determination (OD) software, performed rigorously and its results were used for planning the AMF and STAQ strategies with a greater degree of accuracy. This paper reports the results of evaluation of the performance of the apogee-motor firings employed to place the satellite in its intended position where it is collocated with INSAT-1D satellite. The orbit of the satellite had to be determined continuously at each stage of the initial phase of the mission at a brisk pace and this study shows that the ODS provided consistent results to meet the stringent requirements of the mission operations. At each stage of the mission the orbit was determined using tracking data obtained over varying periods of time. The orbit solutions obtained from short arc OD's are compared with that obtained using the longest arc OD of each stage of the initial phase of the mission. The results of this study have been tabled in this paper. The performance of the ODS in calibrating the ARIANE-4 launch vehicle has been analyzed. A comparison of the orbit elements obtained from the mission operational ODS with the injection parameters provided by CNES, Centre Spatial Guyanais has been made in this paper which shows that the satellite was injected well within the 1 dispersions quoted by ARIANE-SPACE. A comparison has also been shown between the determined transfer orbit elements with pre-launch nominal orbit elements. For the initial phase of this mission ranging support was provided by Hassan earth station at India and INMARSAT network of stations at LakeCowichan (Canada), Fucino (Italy) and Beijing (China). The performance of the tracking systems employed by these stations has been studied. The quality of tracking data obtained from these stations has also been assessed.
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 satellite'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
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 satellite’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
Viking satellite orbit determination
NASA Technical Reports Server (NTRS)
Hildebrand, C. E.; Christensen, E. J.; Boggs, D. H.; Born, G. H.; Hokikian, H.; Jordan, J. F.; Howard, W. B.
1977-01-01
During the summer of 1976, the two Viking spacecraft, each consisting of an orbiter-lander combination, were inserted into orbit about Mars. The paper describes the experiences of the Viking Satellite Orbit Determination Team in determining Mars centered ephemerides of the orbiters and positions of the landers from the two-way Doppler and range data, and synthesizes the different phases of the navigation plan which involves pre-flight modeling and error analysis for all Viking navigation functions from launch through landing. The problem of initial orbit convergence is solved by using DPODP's (Double Precision Orbit Determination Program) square-root batch data filter, and gravity models for both Viking I and Viking II were produced from the combination of short arc estimates. The importance of synchronous Viking orbits (with the rotational period of Mars) is stressed and future extended missions of the spacecraft are outlined.
Roger C. Hart; Kathy R. Hartman; Dipak H. Oza
The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center is currently developing the capability to use the Global Positioning System (GPS) to provide high-accuracy attitude, orbit, and time autonomously onboard NASA spacecraft. NASA's Small Satellite Technology Initiative Lewis spacecraft will host the GPS Attitude Determination Flyer (GADFLY) experiment. The primary objective of GADFLY is to demonstrate the use
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)
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.
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.
Preliminary orbit determination
L. G. Taff; B. Belkin; G. Schweiter; K. Sommar
1989-01-01
The problem of intercontinental ballistic missile reentry vehicle orbit determination from incomplete passive (i.e., angles-only) or active (i.e., distance plus angles) information is presented. Launch location and time and impact location and time to complete the computation of an orbital element set are utilized. In addition, it is possible to construct a good estimate of the variance of the key
Galileo Jupiter approach orbit determination
NASA Technical Reports Server (NTRS)
Miller, J. K.; Nicholson, F. T.
1984-01-01
Orbit determination characteristics of the Jupiter approach phase of the Galileo mission are described. Predicted orbit determination performance is given for the various mission events that occur during Jupiter approach. These mission events include delivery of an atmospheric entry probe, acquisition of probe science data by the Galileo orbiter for relay to earth, delivery of an orbiter to a close encounter of the Galilean satellite Io, and insertion of the orbiter into orbit about Jupiter. The orbit determination strategy and resulting accuracies are discussed for the data types which include Doppler, range, optical imaging of Io, and a new Very Long Baseline Interferometry (VLBI) data type called Differential One-Way Range (DOR).
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.
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.
Lunar Prospector Orbit Determination Results
NASA Technical Reports Server (NTRS)
Beckman, Mark; Concha, Marco
1998-01-01
The orbit support for Lunar Prospector (LP) consists of three main areas: (1) cislunar orbit determination, (2) rapid maneuver assessment using Doppler residuals, and (3) routine mapping orbit determination. The cislunar phase consisted of two trajectory correction maneuvers during the translunar cruise followed by three lunar orbit insertion burns. This paper will detail the cislunar orbit determination accuracy and the real-time assessment of the cislunar trajectory correction and lunar orbit insertion maneuvers. The non-spherical gravity model of the Moon is the primary influence on the mapping orbit determination accuracy. During the first two months of the mission, the GLGM-2 lunar potential model was used. After one month in the mapping orbit, a new potential model was developed that incorporated LP Doppler data. This paper will compare and contrast the mapping orbit determination accuracy using these two models. LP orbit support also includes a new enhancement - a web page to disseminate all definitive and predictive trajectory and mission planning information. The web site provides definitive mapping orbit ephemerides including moon latitude and longitude, and four week predictive products including: ephemeris, moon latitude/longitude, earth shadow, moon shadow, and ground station view periods. This paper will discuss the specifics of this web site.
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.
Mars Science Laboratory Orbit Determination
NASA Technical Reports Server (NTRS)
Kruizinga, Gerhard L.; Gustafson, Eric D.; Thompson, Paul F.; Jefferson, David C.; Martin-Mur, Tomas J.; Mottinger, Neil A.; Pelletier, Frederic J.; Ryne, Mark S.
2012-01-01
This paper describes the orbit determination process, results and filter strategies used by the Mars Science Laboratory Navigation Team during cruise from Earth to Mars. The new atmospheric entry guidance system resulted in an orbit determination paradigm shift during final approach when compared to previous Mars lander missions. The evolving orbit determination filter strategies during cruise are presented. Furthermore, results of calibration activities of dynamical models are presented. The atmospheric entry interface trajectory knowledge was significantly better than the original requirements, which enabled the very precise landing in Gale Crater.
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.
Choosing the Initial LISA Orbital Configuration
NASA Astrophysics Data System (ADS)
Jani, Karan; Finn, Lee Samuel; Benacquista, Mathew
2010-02-01
The Laser Interferometer Space Antenna (LISA) mission proposes to detect gravitational radiation by synthesizing one or more interferometric gravitational wave detectors from fringe velocity measurements generated by chances in the light travel time between three spacecraft in a special set of drag-free, circumsolar orbits. Once the spacecraft are set in their orbits the orientation of the LISA interferometers at any further time is fixed by the Kepler Laws and the initial orientation of the spacecraft constellation. The initial orientation does not affect those locations on the sky where LISA has greatest sensitivity to gravitational waves; however, it does affect those locations where nulls in the LISA response to gravitational waves fall. By artful choice of the LISA initial orientation we can thus choose to optimize LISA's sensitivity to sources or groups of sources whose location (eg., the galactic center or plane, nearby globular cluster, etc.) may be known in advance. )
Calibration effects on orbit determination
NASA Technical Reports Server (NTRS)
Madrid, G. A.; Winn, F. B.; Zielenbach, J. W.; Yip, K. B.
1974-01-01
The effects of charged particle and tropospheric calibrations on the orbit determination (OD) process are analyzed. The calibration process consisted of correcting the Doppler observables for the media effects. Calibrated and uncalibrated Doppler data sets were used to obtain OD results for past missions as well as Mariner Mars 1971. Comparisons of these Doppler reductions show the significance of the calibrations. For the MM'71 mission, the media calibrations proved themselves effective in diminishing the overall B-plane error and reducing the Doppler residual signatures.
Nozomi Cis-Lunar Phase Orbit Determination
NASA Technical Reports Server (NTRS)
Ryne, Mark; Criddle, Kevin
2000-01-01
Japan's Institute of Space and Astronautical Science (ISAS) launched Nozomi, its first mission to the planet Mars using the newly developed M-V launch vehicle on July 3, 1998. Scientific objectives of the mission are to study the structure and dynamics of the Martian upper atmosphere and its interaction with the solar wind. Nozomi is a cooperative mission between ISAS and the National Aeronautics and Space Administration (NASA). The NASA contribution includes navigation and tracking services provided by the Jet Propulsion Laboratory (JPL). The spacecraft also serves as an engineering demonstration of basic technology for planetary exploration. One of the new technologies was a unique trajectory, developed by ISAS, which used solar gravitational perturbations at the weak stability boundary as an aid to achieve an Earth-Mars transfer orbit. This trajectory saves approximately 120 m/s of Delta V compared to direct hyperbolic insertion and is considered an enabling technology for the mission. Nozomi was the first spacecraft to employ this trajectory and provided on-orbit validation of the technique. The trajectory was achieved by initially placing the spacecraft in a highly elliptical cis-lunar phasing orbit. Six maneuvers were performed during this period to correct injection errors and target an outbound lunar swingby in September 1998. The gravity assist from the lunar swingby raised apogee to the vicinity of the weak stability boundary. After three more targeting maneuvers, Nozomi performed an inbound lunar swingby followed immediately by a powered Earth swingby in late December 1998. A 420 m/s Trans Mars Insertion (TMI) burn at the final Earth periapsis was intended to place the spacecraft on a heliocentric trajectory leading to Mars orbit insertion in October 1999. Orbit determination for Nozomi is performed in parallel by both ISAS and the Multi-Mission Navigation (MMNAV) group at JPL. This was an advantage for the mission because each group would generate solutions based on data collected from their respective tracking networks. Spacecraft events, such as sequence uplinks and maneuvers, were generally scheduled during passes at the Usuda tracking station in Japan. As a result, maneuver design and reconstruction was derived from MMNAV solutions based on JPL tracking data obtained immediately prior to or following maneuvers. Data was also exchanged between ISAS and MMNAV so orbit determination could be performed on joint data sets in support of critical targeting late in the cis-lunar phase. In this paper, information regarding the MMNAV orbit determination effort for the first six months of the mission is presented. The spacecraft trajectory is characterized first, followed by a discussion of the orbit determination estimation procedure and models. Results from selected orbit solutions are presented and compared against reconstructed trajectories. One area of emphasis in this paper is orbit determination in the vicinity of the weak stability boundary. Precise navigation was necessary to target the second lunar swingby and the powered Earth swingby. Delivery accuracy of 150 m was required for these critical encounters, but a number of factors contributed to the general degradation of orbit determination accuracy. This included the fact that the spacecraft was at apogee, at a range of 1.7 million km and moving at less than I km/sec perpendicular to the line of sight. Nozomi was also close to zero degrees declination where there are known limitations on orbit determination performance. Finally, S-band tracking data was acquired through the Nozomi backup low gain antenna. This antenna is offset from the axis of this spin stabilized spacecraft and superimposed large signatures in the Doppler and range data. These difficulties were overcome by combining long data arcs, spanning several maneuvers, with a high fidelity solar pressure model. The model included a physically accurate representation of the spacecraft structure and a high time resolution orientation model. Observation modeling included the removal of the spin indu
Advances in SLR Orbit Determination
NASA Astrophysics Data System (ADS)
Kolenkiewicz, R.; Smith, D. E.; Dunn, P. J.; Torrence, M. H.
2002-12-01
Since the LAGEOS II satellite was launched from the space shuttle in October 1992, the Global Laser Tracking Network has provided an extensive set of Satellite Laser Ranging (SLR) data. These data have been used to supplement the 26 years of SLR data obtained from LAGEOS I satellite. The contrasting behavior of small, unmodeled along-track accelerations in the orbits of each satellite can be used to help explain the sources of these perturbations. LAGEOS II's inclination provides gravity and tidal sensitivity to improve on the advances from LAGEOS I in a shorter time, as LAGEOS II's nodal precession period is about one half that of LAGEOS I's three year period. The geometry added by LAGEOS II has been found to improve station positioning accuracy to better than one centimeter in latitude, longitude and height. This is an asset for the determination of the regional deformation in the vicinity of the SLR sites. The improved SLR global reference frame can now be used to accurately define orbit and station positions for scientific applications and the support of altimeter missions.
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
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.
Precise Orbit Determination for FORMOSAT-3/COSMIC and Gravity Application
NASA Astrophysics Data System (ADS)
Hwang, C.; Tseng, T.; Lin, T.; Fu, C.; Svehla, D.
2006-12-01
The orbits of FORMOSAT-3/COSMIC (FC) are determined using the GPS data at a 5-s sampling rate. Both reduced dynamic and kinematic solutions are employed. GPS data from only one antenna is used. The average RMS orbital differences between the reduced dynamic and the kinematic solutions for FM1 to FM6 is 6 cm in each coordinate component. The RMS orbital difference at the overlapped arc is regarded as the internal orbital accuracy. Typical RMS overlapping differences from the reduced and kinematic solutions are about 5 cm. While the reduced dynamic orbit is smooth and free from outliers in most cases, the kinematic solution yields large orbit errors when the satellite attitude values are large or missing. Expected orbital improvement can be achieved by combining GPS data from the two POD antennae. As an external accuracy assessment, we numerically integrate FC orbits using the initial state vectors from the GPS solutions and modeled perturbing forces acting on FC satellites. The RMS differences between GPS-determined and numerically integrated orbits are about 10 cm. For gravity application, we model the centers of mass of the six FC satellites and their time variations by considering the consumption of fuel and motion of solar panels. We will also present experimental determinations of gravity harmonic coefficients using kinematic orbits of FC in different scenarios. The result of a combination solution of gravity harmonic coefficients using FC and GRACE data will also be presented.
Precise Orbit Determination for FORMOSAT-3/COSMIC and Gravity Application
NASA Astrophysics Data System (ADS)
Hwang, C.; Tseng, T.; Lin, T.; Fu, C.; Svehla, D.
2005-05-01
The orbits of FORMOSAT-3/COSMIC (FC) are determined using the GPS data at a 5-s sampling rate. Both reduced dynamic and kinematic solutions are employed. GPS data from only one antenna is used. The average RMS orbital differences between the reduced dynamic and the kinematic solutions for FM1 to FM6 is 6 cm in each coordinate component. The RMS orbital difference at the overlapped arc is regarded as the internal orbital accuracy. Typical RMS overlapping differences from the reduced and kinematic solutions are about 5 cm. While the reduced dynamic orbit is smooth and free from outliers in most cases, the kinematic solution yields large orbit errors when the satellite attitude values are large or missing. Expected orbital improvement can be achieved by combining GPS data from the two POD antennae. As an external accuracy assessment, we numerically integrate FC orbits using the initial state vectors from the GPS solutions and modeled perturbing forces acting on FC satellites. The RMS differences between GPS-determined and numerically integrated orbits are about 10 cm. For gravity application, we model the centers of mass of the six FC satellites and their time variations by considering the consumption of fuel and motion of solar panels. We will also present experimental determinations of gravity harmonic coefficients using kinematic orbits of FC in different scenarios. The result of a combination solution of gravity harmonic coefficients using FC and GRACE data will also be presented.
Orbit determination and control for the European Student Moon Orbiter
NASA Astrophysics Data System (ADS)
Zuiani, Federico; Gibbings, Alison; Vetrisano, Massimo; Rizzi, Francesco; Martinez, Cesar; Vasile, Massimiliano
2012-10-01
This paper presents the preliminary navigation and orbit determination analyses for the European Student Moon Orbiter. The severe constraint on the total mission ?v and the all-day piggy-back launch requirement imposed by the limited available budget, led to the choice of using a low-energy transfer, more specifically a Weak Stability Boundary one, with a capture into an elliptic orbit around the Moon. A particular navigation strategy was devised to ensure capture and fulfil the requirement for the uncontrolled orbit stability at the Moon. This paper presents a simulation of the orbit determination process, based on an extended Kalman filter, and the navigation strategy applied to the baseline transfer of the 2011-2012 window. The navigation strategy optimally allocates multiple Trajectory Correction Manoeuvres to target a so-called capture corridor. The capture corridor is defined, at each point along the transfer, by back-propagating the set of perturbed states at the Moon that provides an acceptable lifetime of the lunar orbit.
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.
TDRS orbit determination by radio interferometry
NASA Astrophysics Data System (ADS)
Pavloff, Michael S.
1994-11-01
In support of a NASA study on the application of radio interferometry to satellite orbit determination, MITRE developed a simulation tool for assessing interferometry tracking accuracy. The Orbit Determination Accuracy Estimator (ODAE) models the general batch maximum likelihood orbit determination algorithms of the Goddard Trajectory Determination System (GTDS) with the group and phase delay measurements from radio interferometry. ODAE models the statistical properties of tracking error sources, including inherent observable imprecision, atmospheric delays, clock offsets, station location uncertainty, and measurement biases, and through Monte Carlo simulation, ODAE calculates the statistical properties of errors in the predicted satellites state vector. This paper presents results from ODAE application to orbit determination of the Tracking and Data Relay Satellite (TDRS) by radio interferometry. Conclusions about optimal ground station locations for interferometric tracking of TDRS are presented, along with a discussion of operational advantages of radio interferometry.
TDRS orbit determination by radio interferometry
NASA Technical Reports Server (NTRS)
Pavloff, Michael S.
1994-01-01
In support of a NASA study on the application of radio interferometry to satellite orbit determination, MITRE developed a simulation tool for assessing interferometry tracking accuracy. The Orbit Determination Accuracy Estimator (ODAE) models the general batch maximum likelihood orbit determination algorithms of the Goddard Trajectory Determination System (GTDS) with the group and phase delay measurements from radio interferometry. ODAE models the statistical properties of tracking error sources, including inherent observable imprecision, atmospheric delays, clock offsets, station location uncertainty, and measurement biases, and through Monte Carlo simulation, ODAE calculates the statistical properties of errors in the predicted satellites state vector. This paper presents results from ODAE application to orbit determination of the Tracking and Data Relay Satellite (TDRS) by radio interferometry. Conclusions about optimal ground station locations for interferometric tracking of TDRS are presented, along with a discussion of operational advantages of radio interferometry.
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 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.
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 ac
A Robust Method of Preliminary Orbit Determination
NASA Astrophysics Data System (ADS)
Wang, X.
2013-05-01
The preliminary orbit determination with optical angular measurements plays an important role in the survey of space object.The classical method of orbit computing, based on least square error estimation is not robust while the outliers occur in the observation. A robust method is proposed by employing the least absolute deviation estimation. The method reduces the problem of orbit computing to a linear programming problem, and gives the variance of the estimation with bootstrap method. Numerical check shows that the method is effective as well as robust, and has a high breakdown point.
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.
Comparison of four European orbit determination systems
NASA Astrophysics Data System (ADS)
Alby, F.; Bianco, G.; Fourcade, J.; Gill, E.; Kirschner, M.; Luceri, V.; Mesnard, R.; Montenbruck, O.; Schneller, M.; Schoemaekers, J.
The objective of this paper is to present the activities and the main conclusions of the IOI (In-Orbit Infrastructure) Flight Dynamics Working Group. This IOIFDWG is an inter-agency working group composed of flight dynamics experts coming from the following space agencies: ASI, DLR, ESA and CNES. These agencies are in charge of elements of the future European orbital infrastructure such as the space station, the launcher, the vehicles, relay satellites or the ground control center. During the operations they will have to exchange flight dynamics data; therefore the main objective of the group was to analyze the compatibility of the orbit determination systems at the participating agencies. In order to identify the origin of possible discrepancies in the complex orbit determination process, the comparisons have been performed in three successive steps: first the orbit propagation software modules have been compared, then the tracking measurements modules have been compared and, finally the orbit estimation software modules themselves. After a short description of the software systems used at each agency, with their main features, this paper presents, for each step, the description of the tests and the main results. Then, the significant differences are discussed and in particular, the major contribution due to the atmospheric drag modeling is shown; finally, the main conclusions are summarized and recommendations for the exchange of state vectors are proposed including a list of data to be harmonized between agencies.
New determination of the orbit of Miranda
NASA Astrophysics Data System (ADS)
Veillet, C.
1981-05-01
Determinations of the orbital elements of Miranda, the innermost satellite of Uranus, which were made subsequent to the determinations of Whitaker and Greenberg (1973) are reported. The present determination is based on observations made during the 1975 and 1977 oppositions with the 155-cm astrometric reflector at Flagstaff Observatory, and during the successive oppositions of 1977, 1978 and 1979 with the Pic-du-Midi 1-m reflector. Comparison of the observed position angles with those predicted from the orbit of Whitaker and Greenberg reveals a discrepancy increasing with time until 1978, indicating the need for an orbital revision. Calculations of a sinusoidal approximation of the residuals in longitude observed in 1948-49 and since 1972 results in a period close to the circulation period of near commensurability between Ariel, Umbriel and Miranda, which may also be interpreted as the gravitational effect of Ariel and Umbriel on Miranda. Determinations of the orbit of Miranda taking into account these gravitational perturbations confirm the previously determined irregularities in inclination and eccentricity. A mass product of Ariel and Umbriel of 1.10 plus or minus 0.25 x 10 to the -10th is also derived. It is noted that observations are continuing to improve the accuracy of the present results.
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.
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.
GPS-assisted GLONASS orbit determination
D. Kuang; Y. E. Bar-Sever; W. I. Bertiger; K. J. Hurst; J. F. Zumberge
2001-01-01
. ?Using 1 week of data from a network of GPS\\/GLONASS dual-tracking receivers, 15-cm accurate GLONASS orbit determination is\\u000a demonstrated with an approach that combines GPS and GLONASS data. GPS data are used to define the reference frame, synchronize\\u000a receiver clocks and determine troposphere delay for the GLONASS tracking network. GLONASS tracking data are then processed\\u000a separately, with the GPS-defined parameters
NASA Astrophysics Data System (ADS)
Hugentobler, U.; Beutler, G.
2003-07-01
Considerable experience accumulated during the past decade in strategies for processing GPS data from ground-based geodetic receivers. First experience on the use of GPS observations from spaceborne receivers for orbit determination of satellites on low altitude orbits was gained with the launch of TOPEX/POSEIDON ten years ago. The launch of the CHAMP satellite in July 2000 stimulated a number of activities worldwide on improving the strategies and algorithms for orbit determination for Low Earth Orbiters (LEOs) using the GPS. Similar strategies as for ground-based receivers are applied to data from spaceborne GPS receivers to determine high precision orbits. Zero- and double-differencing techniques are applied to obtain kinematic and/or reduced-dynamic orbits with an accuracy which is today at the decimeter level. Further developments in modeling and processing strategies will continuously improve the quality of GPS-derived LEO orbits in the near future. A significant improvement can be expected from fixing double-difference phase ambiguities to integer numbers. Particular studies focus on the impact of a combined processing of LEO and GPS orbits on the quality of orbits and the reference frame realization.
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.
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.
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.
Precision orbit determination software validation experiment
NASA Technical Reports Server (NTRS)
Schutz, B. E.; Tapley, B. D.; Eanes, R. J.; Marsh, J. G.; Williamson, R. G.; Martin, T. V.
1980-01-01
This paper presents the results of an experiment which was designed to ascertain the level of agreement between GEODYN and UTOPIA, two completely independent computer programs used for precision orbit determination, and to identify the sources which limit the agreement. For a limited set of models and a seven-day data set arc length, the altitude components of the ephemeris obtained by the two programs agree at the sub-centimeter level throughout the arc.
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 via adaptive Gaussian swarm optimization
NASA Astrophysics Data System (ADS)
Kiani, Maryam; Pourtakdoust, Seid H.
2015-02-01
Accurate orbit determination (OD) is vital for every space mission. This paper proposes a novel heuristic filter based on adaptive sample-size Gaussian swarm optimization (AGSF). The proposed estimator considers the OD as a stochastic dynamic optimization problem that utilizes a swarm of particles in order to find the best estimation at every time step. One of the key contributions of this paper is the adaptation of the swarm size using a weighted variance approach. The proposed strategy is simulated for a low Earth orbit (LEO) OD problem utilizing geomagnetic field measurements at 700 km altitude. The performance of the proposed AGSF is verified using Monte Carlo simulation whose results are compared with other advanced sample based nonlinear filters. It is demonstrated that the adopted filter achieves about 2.5 km accuracy in position estimation that fulfills the essential requirements of accuracy and convergence time for OD problem.
Information measures for statistical orbit determination
NASA Astrophysics Data System (ADS)
Mashiku, Alinda K.
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 estimates of the state and the statistical information of its region of uncertainty given by the probability density function (PDF). As even collision events with very low probability are important, accurate prediction of collisions require the representation of the full 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. In order to obtain an accurate representation of the PDF that incorporates higher order statistical information, we propose the use of nonlinear estimation methods such as the Particle Filter. 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. Moreover, during longer state propagations, we propose to represent the final state vector as a compressed probability mass function (PMF). Multivariate PDF compressions are computationally costly and could potentially be numerically intractable. We tackle this issue by decorrelating the nonlinear multivariate state PMFs using an improved nonlinear factor analysis (NFA) that uses a multilayer perceptron (MLP) network to model the state nonlinearities and obtain the sources that also incorporates the Fast Independent Component Analysis (FastICA [a faster computational method for ICA]) to obtain the independent and decorrelated states. Methods such as the Principal Component Analysis (PCA) are based on utilizing moments that only incorporate the second order statistics, hence will not suffice in maintaining maximum information content. On the other hand, the Independent Component Analysis (ICA) is a non-Gaussian decorrelator that is based on a linear mapping scheme, that does not incorporate the non-linear information. The PDF compressions are achieved by implementing the fast-Fourier Transform (FFT) and the wavelet transform (WT) to construct a smaller subset of data for data allocation and transmission cost reduction. The accuracy of tracking the space objects as well as reduced costs will help increase the capability of tracking the increased number of space objects. We use statistical information measures such as the Kolmogorov-Smirnov (K-S) test and the Kullback-Leibler Divergence (KLD) metric to quantify the accuracy of the reconstructed state vector and the cost reduction is measured by the number of terms required to represent the states. A performance plot illuminates the performances of the transforms over a range of compression rates. Simulations are performed on real and simulated data to demonstrate the approach for this work.
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.
49 CFR 7.31 - Initial determinations.
Code of Federal Regulations, 2011 CFR
2011-10-01
...Initial determinations. An initial determination whether to release a record requested pursuant to subpart C of this...component using multitrack processing may provide requesters in its slower track(s) with an opportunity to limit the scope of...
32 CFR 518.16 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-07-01
...2010-07-01 true Initial determinations. 518.16 Section...518.16 Initial determinations. (a) Initial denial...Each IDA will act on direct and referred...dissatisfied with adverse determinations. It is crucial to...the IDA by name and position in the written...
Orbit determination singularities in the Doppler tracking of a planetary orbiter
NASA Technical Reports Server (NTRS)
Wood, L. J.
1985-01-01
On a number of occasions, spacecraft launched by the U.S. have been placed into orbit about the moon, Venus, or Mars. It is pointed out that, in particular, in planetary orbiter missions two-way coherent Doppler data have provided the principal data type for orbit determination applications. The present investigation is concerned with the problem of orbit determination on the basis of Doppler tracking data in the case of a spacecraft in orbit about a natural body other than the earth or the sun. Attention is given to Doppler shift associated with a planetary orbiter, orbit determination using a zeroth-order model for the Doppler shift, and orbit determination using a first-order model for the Doppler shift.
Toward decimeter Topex orbit determination using GPS
NASA Technical Reports Server (NTRS)
Wu, Sien-Chong; Yunck, Thomas P.; Hajj, George A.
1990-01-01
Several practical aspects of precision GPS-based Topex orbit determination are investigated. Multipath signals contaminating Topex pseudorange data are greatly reduced by placing the GPS antenna on a conducting backplate consisting of concentric choke rings to attenuate signals coming in from the Topex horizon and below, and by elevating it on a boom to keep it well above all reflecting surfaces. A proper GPS antenna cutoff view angle is chosen so that a sufficient number of GPS satellites with good geometry are in view while reception of reflected signals is minimized. The geometrical strength of the tracking data is optimized by properly selecting GPS satellites to be observed so as to provide data with moderate continuity, low PDOP, and common visibility with ground tracking sites. The tracking performance is greatly enhanced when three complementary sites are added to the minimum ground tracking network consisting of the three NASA DSN sites.
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.
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.
Voyager 1 and Voyager 2 Saturn encounter orbit determination
NASA Technical Reports Server (NTRS)
Campbell, J. K.; Jacobson, R. A.; Riedel, J. E.; Synnott, S. P.; Taylor, A. H.
1982-01-01
This paper contains quantitative results and conclusions from the Saturn approach orbit determination for Voyagers 1 and 2. The major topics covered include an overview of the navigation-related requirements and a review of the salient orbit determination results obtained. Special attention is paid to the use of combined spacecraft-based optical observations and earth-based radiometric observations to achieve accurate orbit determination during the Saturn encounter approach phase.
Optimal solutions of unobservable orbit determination problems
NASA Astrophysics Data System (ADS)
Cicci, David A.; Tapley, Byron D.
1988-12-01
The method of data augmentation, in the form ofa priori covariance information on the reference solution, as a means to overcome the effects of ill-conditioning in orbit determination problems has been investigated. Specifically, for the case when ill-conditioning results from parameter non-observability and an appropriatea priori covariance is unknown, methods by which thea priori covariance is optimally chosen are presented. In problems where an inaccuratea priori covariance is provided, the optimal weighting of this data set is obtained. The feasibility of these ‘ridge-type’ solution methods is demonstrated by their application to a non-observable gravity field recovery simulation. In the simulation, both ‘ridge-type’ and conventional solutions are compared. Substantial improvement in the accuracy of the conventional solution is realized by the use of these ridge-type solution methods. The solution techniques presented in this study are applicable to observable, but ill-conditioned problems as well as the unobservable problems directly addressed. For the case of observable problems, the ridge-type solutions provide an improvement in the accuracy of the ordinary least squares solutions.
EURECA 11 months in orbit: Initial post flight investigation results
NASA Technical Reports Server (NTRS)
Dover, Alan; Aceti, Roberto; Drolshagen, Gerhard
1995-01-01
This paper gives a brief overview of the European free flying spacecraft 'EURECA' and the initial post flight investigations following its retrieval in June 1993. EURECA was in low earth orbit for 11 months commencing in August 1992, and is the first spacecraft to be retrieved and returned to Earth since the recovery of LDEF. The primary mission objective of EURECA was the investigation of materials and fluids in a very low micro-gravity environment. In addition other experiments were conducted in space science, technology and space environment disciplines. The European Space Agency (ESA) has taken the initiative in conducting a detailed post-flight investigation to ensure the full exploitation of this unique opportunity.
Jason-2 Precise Orbit Determination : current status and future improvements
NASA Astrophysics Data System (ADS)
Cerri, Luca
The JASON-2 satellite was launched on June 20, 2008 to continue the series of spaceborne radar altimeter missions initiated with TOPEX-POESEIDON in 1992 and continued by its follow-on, JASON-1, starting in 2002. From the very beginning, Precise Orbit Determination (POD) has been a key component of the success of these satellite altimeter missions. In order to meet the 1.5 cm radial accuracy required for the operational precise orbits included in the Geophysical Data Record (GDR), both JASON satellites are equipped with three state-of-the-art track-ing systems: Doppler Orbitography and Radiopositioning Integrated by Satellite (DORIS) , Satellite Laser Ranging (SLR), and Global Positioning System (GPS). Over the past 10 years, several improvements in the quality as well as in the spatial and the temporal coverage of these tracking data, together with the enhancements in models and parameterization techniques, have made it possible to achieve the "1-cm" goal. Today, JASON-2 orbits computed by various research groups compare at the sub-centimeter level in terms of radial RMS over a few days. A significant effort is now committed toward reducing the latency of the precise orbit products while maintaining a high level of accuracy and operational robustness. Most important, as the altimeter data set now spans over almost two decades, scientists are able to reveal small climate signals such as a 3mm/year rise of the global Mean Sea Level (MSL) as well as interannual fluctuations of few mm amplitude. To maintain this level of performance, new requirements on the long term drifts of all components of the measurement system of future altimeter missions are needed. In this context, the stability of the radial orbit error properties over several years is increasingly important. Typically, global RMS values can cover systematic variations that have a particular spatial and temporal coherence, and which are of particular interest for the altimeter data analysts. In particular, errors in the DORIS, SLR and GPS realizations of the terrestrial reference system in which orbit solutions are computed, mismodeled surface forces and temporal variations of the gravity field are fundamental contributors to the orbit error budget at the global and local scales. Two years after launch, this talk addresses these topics with an overview of JASON-2 POD performance, both in term of short term and long term accuracy, outlining past progress and prospects for future improvements.
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.
TAOS Orbit Determination Results lJsing Global Positioning Satellites
Joseph R; Bobby G. Williams; Peter J. Wolff
Orbit determination results for the Air Force Phillips Laboratory's Technology for Autonomous Operational Survivability (TAOS) satellite using a Rockwell AST V Global Positioning System (GPS) receiver are presented in this paper. Under a cooperative effort, GPS orbit determination technology developed at Jet Propulsion Laboratory (JPL) has been transferred to the U.S. Air Force. JPL's post processing differential GPS software MIRAGE
The determination of the satellite orbit of Mariner 9.
NASA Technical Reports Server (NTRS)
Born, G. H.; Christensen, E. J.; Ferrari, A. J.; Jordan, J. F.; Reinbold, S. J.
1972-01-01
This paper presents a comprehensive analysis of the Mars orbital phase of the Mariner 9 trajectory as determined from Earth based radio data. Both the method and accuracy of the orbit determination process are reviewed. Analysis is presented to show the effects of Mars gravity model and node in the plane of the sky errors on the accuracy of orbit determination. In addition the long term evolution of the orbit from insertion to date is presented, and is decomposed into effects from the Mars gravity field, n-body perturbations, and solar radiation pressure. Since the orbit period is nearly commensurable with the Mars rotational period, the orbit experiences significant resonance perturbations. The primary perturbation is in-track with a maximum amplitude of 1000 km and a wavelength of 39 revolutions.
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.
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.
Phenomenological Determination of the Orbital Angular Momentum
Ramsey, Gordon P. [Physics Department, Loyola University, Chicago, IL 60626 (United States) and High Energy Physics Division, Argonne National Lab, Argonne, IL 60439 (United States)
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.
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.
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.
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.
Orbit determination for low-thrust spacecraft: Concepts and analysis
NASA Technical Reports Server (NTRS)
Mcdanell, J. P.
1973-01-01
Earth-based orbit determination capability for SEP spacecraft in multistation tracking and in thrust subsystem error modeling is described. Five different tracking strategies are applied to a 15 day segment of an Encke rendezvous mission. Both optimal and suboptimal orbit determination performance are determined for a wide range of process noise parameter values. The multi-station tracking techniques are found to be extremely effective, reducing orbit determination errors by orders of magnitude over that obtained with conventional single-station tracking. Explicitly differenced multistation data (QVLBI) is found to be least sensitive to gross modeling errors, but if a reasonably good process noise model is available, explicit differencing is not required.
20 CFR 320.5 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-04-01
... REGULATIONS UNDER THE RAILROAD UNEMPLOYMENT INSURANCE ACT INITIAL DETERMINATIONS UNDER THE RAILROAD UNEMPLOYMENT INSURANCE ACT AND REVIEWS OF AND...made with respect to each claim for unemployment or sickness benefits...
32 CFR 300.8 - Initial determinations.
Code of Federal Regulations, 2014 CFR
2014-07-01
...of Defense (Continued) OFFICE OF THE SECRETARY OF DEFENSE (CONTINUED) FREEDOM OF INFORMATION ACT PROGRAM DEFENSE LOGISTICS AGENCY FREEDOM OF INFORMATION ACT PROGRAM FOIA Request Processing § 300.8 Initial determinations....
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2013 CFR
2013-07-01
...ORDER 13526 Action on Challenges § 1907.24. Initial determination. (a) Formal challenges shall be directed to the CIA Information and Privacy Coordinator (Coordinator) who shall promptly forward the challenge to the C/CMCG for action....
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2012 CFR
2012-07-01
...ORDER 13526 Action on Challenges § 1907.24. Initial determination. (a) Formal challenges shall be directed to the CIA Information and Privacy Coordinator (Coordinator) who shall promptly forward the challenge to the C/CMCG for action....
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2014 CFR
2014-07-01
...ORDER 13526 Action on Challenges § 1907.24. Initial determination. (a) Formal challenges shall be directed to the CIA Information and Privacy Coordinator (Coordinator) who shall promptly forward the challenge to the C/CMCG for action....
Initial observations from the Lunar Orbiter Laser Altimeter (LOLA)
Smith, David Edmund
As of June 19, 2010, the Lunar Orbiter Laser Altimeter, an instrument on the Lunar Reconnaissance Orbiter, has collected over 2.0 × 109 measurements of elevation that collectively represent the highest resolution global ...
An exact solution to determination of an open orbit
Hideki Asada
2006-10-23
We present an exact solution of the equations for orbit determination of a two body system in a hyperbolic or parabolic motion. In solving this problem, we extend the method employed by Asada, Akasaka and Kasai (AAK) for a binary system in an elliptic orbit. The solutions applicable to each of elliptic, hyperbolic and parabolic orbits are obtained by the new approach, and they are all expressed in an explicit form, remarkably, only in terms of elementary functions. We show also that the solutions for an open orbit are recovered by making a suitable transformation of the AAK solution for an elliptic case.
Precision orbit determination at the NASA Goddard Space Flight Center
NASA Technical Reports Server (NTRS)
Putney, B.; Kolenkiewicz, R.; Smith, D.; Dunn, P.; Torrence, M. H.
1990-01-01
This paper describes the GEODYN computer program developed by the Geodynamics Branch at the NASA Goddard Space Flight Center and outlines the procedure for accurate satellite orbit and tracking-data analyses. The capabilities of the program allow the development of gravity fields as large as 90 by 90, and a complete modeling of tidal parameters. It is also feasible to numerically integrate a continuous orbit of a satellite such as Lageos for up to 12 years. The evolution of the orbit can be studied, and, by comparison with locally determined orbits, force model improvements can be made. The GEODYN flow diagrams are presented.
NASA Astrophysics Data System (ADS)
Tukaram Aghav, Sandip; Achyut Gangal, Shashikala
2014-06-01
In this paper, the main work is focused on designing and simplifying the orbit determination algorithm which will be used for Low Earth Orbit (LEO) navigation. The various data processing algorithms, state estimation algorithms and modeling forces were studied in detail, and simplified algorithm is selected to reduce hardware burden and computational cost. This is done by using raw navigation solution provided by GPS Navigation sensor. A fixed step-size Runge-Kutta 4th order numerical integration method is selected for orbit propagation. Both, the least square and Extended Kalman Filter (EKF) orbit estimation algorithms are developed and the results of the same are compared with each other. EKF algorithm converges faster than least square algorithm. EKF algorithm satisfies the criterions of low computation burden which is required for autonomous orbit determination. Simple static force models also feasible to reduce the hardware burden and computational cost.
The GPS based precision orbit determination experiment on TOPEX
NASA Technical Reports Server (NTRS)
Melbourne, William G.; Davis, Edgar S.; Yunck, Thomas P.
1988-01-01
The objectives of the GPS-based precision orbit determination (POD) experiment on TOPEX are discussed. Problems facing this experiment include the careful design of all network receivers to control uncalibrated systematic group-delay biases and delay variations between channels, and the careful design of both the GPS-antenna-TOPEX satellite interface and the ground antennas to mimimize multipath. Questions of reference frames, geoid recovery, and the application of innovative orbit determination strategies must also be addressed.
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.
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.
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.
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 SAAMER’s location, the system is ideal for providing crucial data to continuously study the South Toroidal and South Apex sporadic meteoroid apparent sources.
Cassini Orbit Determination Results: January 2006 - End of Prime Mission
NASA Technical Reports Server (NTRS)
Antreasian, P. G.; Ardalan, S. M.; 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.; Thompson, P. F.; Vaughan, A. T.
2008-01-01
After the forty-fifth flyby of Titan, the Cassini spacecraft has successfully completed the planned four-year prime mission tour of the Saturnian system. This paper reports on the orbit determination performance of the Cassini spacecraft over two years spanning 2006 - 2008. In this time span, Cassini's orbit progressed through the magnetotail and pi-transfer phases of the mission. Thirty-four accurate close encounters of Titan, one close flyby of Iapetus and one 50 km flyby of Enceladus were performed during this period. The Iapetus and Enceladus flybys were especially challenging and so the orbit determination supporting these encounters will be discussed in more detail. This paper will show that in most cases orbit determination has exceeded the navigation requirements for targeting flybys and predicting science instrument pointing during these encounters.
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.
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.
Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation
NASA Technical Reports Server (NTRS)
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.
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.
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.
GOCE: precise orbit determination for the entire mission
NASA Astrophysics Data System (ADS)
Bock, Heike; Jäggi, Adrian; Beutler, Gerhard; Meyer, Ulrich
2014-07-01
The Gravity field and steady-state Ocean Circulation Explorer (GOCE) was the first Earth explorer core mission of the European Space Agency. It was launched on March 17, 2009 into a Sun-synchronous dusk-dawn orbit and re-entered into the Earth's atmosphere on November 11, 2013. The satellite altitude was between 255 and 225 km for the measurement phases. The European GOCE Gravity consortium is responsible for the Level 1b to Level 2 data processing in the frame of the GOCE High-level processing facility (HPF). The Precise Science Orbit (PSO) is one Level 2 product, which was produced under the responsibility of the Astronomical Institute of the University of Bern within the HPF. This PSO product has been continuously delivered during the entire mission. Regular checks guaranteed a high consistency and quality of the orbits. A correlation between solar activity, GPS data availability and quality of the orbits was found. The accuracy of the kinematic orbit primarily suffers from this. Improvements in modeling the range corrections at the retro-reflector array for the SLR measurements were made and implemented in the independent SLR validation for the GOCE PSO products. The satellite laser ranging (SLR) validation finally states an orbit accuracy of 2.42 cm for the kinematic and 1.84 cm for the reduced-dynamic orbits over the entire mission. The common-mode accelerations from the GOCE gradiometer were not used for the official PSO product, but in addition to the operational HPF work a study was performed to investigate to which extent common-mode accelerations improve the reduced-dynamic orbit determination results. The accelerometer data may be used to derive realistic constraints for the empirical accelerations estimated for the reduced-dynamic orbit determination, which already improves the orbit quality. On top of that the accelerometer data may further improve the orbit quality if realistic constraints and state-of-the-art background models such as gravity field and ocean tide models are used for the reduced-dynamic orbit determination.
Evaluation of semiempirical atmospheric density models for orbit determination applications
NASA Astrophysics Data System (ADS)
Cox, C. M.; Feiertag, R. J.; Oza, D. H.; Doll, C. E.
1994-05-01
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.
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.
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.
NASA Technical Reports Server (NTRS)
Folkner, W. M.; Border, J. S.; Nandi, S.; Zukor, K. S.
1993-01-01
A new radio metric positioning technique has demonstrated improved orbit determination accuracy for the Magellan and Pioneer Venus Orbiter orbiters. The new technique, known as Same-Beam Interferometry (SBI), is applicable to the positioning of multiple planetary rovers, landers, and orbiters which may simultaneously be observed in the same beamwidth of Earth-based radio antennas. Measurements of carrier phase are differenced between spacecraft and between receiving stations to determine the plane-of-sky components of the separation vector(s) between the spacecraft. The SBI measurements complement the information contained in line-of-sight Doppler measurements, leading to improved orbit determination accuracy. Orbit determination solutions have been obtained for a number of 48-hour data arcs using combinations of Doppler, differenced-Doppler, and SBI data acquired in the spring of 1991. Orbit determination accuracy is assessed by comparing orbit solutions from adjacent data arcs. The orbit solution differences are shown to agree with expected orbit determination uncertainties. The results from this demonstration show that the orbit determination accuracy for Magellan obtained by using Doppler plus SBI data is better than the accuracy achieved using Doppler plus differenced-Doppler by a factor of four and better than the accuracy achieved using only Doppler by a factor of eighteen. The orbit determination accuracy for Pioneer Venus Orbiter using Doppler plus SBI data is better than the accuracy using only Doppler data by 30 percent.
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.
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 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.
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.
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.
42 CFR 476.83 - Initial denial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
...2010-10-01 false Initial denial determinations. 476...FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH... § 476.83 Initial denial determinations. A...that the health care services furnished or proposed...care, is an initial denial determination and...
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.
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.
Modeling GPS satellite attitude variation for precise orbit determination
D. Kuang; H. J. Rim; B. E. Schutz; P. A. M. Abusali
1996-01-01
High precision geodetic application of the Global Positioning System (GPS) require highly precise ephemerides of the GPS satellites. An accurate model for the non-gravitational forces on the GPS satellites is a key to high quality GPS orbit determination, especially in long arcs. In this paper the effect of the satellite solar panel orientation error is investigated. These effects are approximated
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.
Implementation of a low-cost, commercial orbit determination system
NASA Astrophysics Data System (ADS)
Corrigan, Jim
1994-05-01
This paper describes the implementation and potential applications of a workstation-based orbit determination system developed by Storm Integration, Inc. called the Precision Orbit Determination System (PODS). PODS is offered as a layered product to the commercially-available Satellite Tool Kit (STK) produced by Analytical Graphics, Inc. PODS also incorporates the Workstation/Precision Orbit Determination (WS/POD) product offered by Van Martin System, Inc. The STK graphical user interface is used to access and invoke the PODS capabilities and to display the results. WS/POD is used to compute a best-fit solution to user-supplied tracking data. PODS provides the capability to simultaneously estimate the orbits of up to 99 satellites based on a wide variety of observation types including angles, range, range rate, and Global Positioning System (GPS) data. PODS can also estimate ground facility locations, Earth geopotential model coefficients, solar pressure and atmospheric drag parameters, and observation data biases. All determined data is automatically incorporated into the STK data base, which allows storage, manipulation and export of the data to other applications. PODS is offered in three levels: Standard, Basic GPS and Extended GPS. Standard allows processing of non-GPS observation types for any number of vehicles and facilities. Basic GPS adds processing of GPS pseudo-ranging data to the Standard capabilities. Extended GPS adds the ability to process GPS carrier phase data.
Implementation of a low-cost, commercial orbit determination system
NASA Technical Reports Server (NTRS)
Corrigan, Jim
1994-01-01
This paper describes the implementation and potential applications of a workstation-based orbit determination system developed by Storm Integration, Inc. called the Precision Orbit Determination System (PODS). PODS is offered as a layered product to the commercially-available Satellite Tool Kit (STK) produced by Analytical Graphics, Inc. PODS also incorporates the Workstation/Precision Orbit Determination (WS/POD) product offered by Van Martin System, Inc. The STK graphical user interface is used to access and invoke the PODS capabilities and to display the results. WS/POD is used to compute a best-fit solution to user-supplied tracking data. PODS provides the capability to simultaneously estimate the orbits of up to 99 satellites based on a wide variety of observation types including angles, range, range rate, and Global Positioning System (GPS) data. PODS can also estimate ground facility locations, Earth geopotential model coefficients, solar pressure and atmospheric drag parameters, and observation data biases. All determined data is automatically incorporated into the STK data base, which allows storage, manipulation and export of the data to other applications. PODS is offered in three levels: Standard, Basic GPS and Extended GPS. Standard allows processing of non-GPS observation types for any number of vehicles and facilities. Basic GPS adds processing of GPS pseudo-ranging data to the Standard capabilities. Extended GPS adds the ability to process GPS carrier phase data.
Orbit determination and control for the AMPTE UK satellite
G. H. Spalding
1986-01-01
The procedures used for orbit determination and control of the AMPTE (Active Magnetospheric Particle Tracer Explorers) UK satellite (UKS) are described and the operational experience is reviewed. In particular, details are given about how the separation between the UKS and the companion Ion Release Module satellite was controlled to meet the scientific requirements of the AMPTE mission, utilising a cold
NASA Astrophysics Data System (ADS)
Syusina, O. M.; Chernitsov, A. M.; Tamarov, V. A.; Baturin, A. P.
2011-07-01
The analysis various systems of initial orbital elements of comet Herschel-Rigollet defined in bases on different sample of observations was given. In spite of slight quantity of first appearance observations the introduction of weighting coefficients and the new rejection algorithm is allowed to define the most precise system of orbital elements with the least value of volume confidence region.
Salyut-7/Kosmos-1686 orbit determination from radar data
NASA Astrophysics Data System (ADS)
Jehn, R.
1991-08-01
During the final reentry phase of the Salyut-7 orbital complex, transmissions of radar data were received weekly (in January) and daily (last seven days). The radar data comprised measurements of slant range, range rate, azimuth, and elevation. The data were processed by an iterative least squares algorithm to derive the state vector and the ballistic coefficient of the space station. The algorithm is explained, and critical areas where straightforward convergence is hampered are highlighted. Methods to solve the problem of convergence are presented together with the solutions of the orbit determination.
Capabilities of a single TDRS to support user orbit determination
NASA Technical Reports Server (NTRS)
Cappellari, J. O., Jr.; Kay, P. Y.; Nicholson, A. M.
1988-01-01
It is shown that the single-TDRS S-band tracking configuration satisfies the navigation certification requirements for operational orbit determination support for the Landsat-5, SMM, SME, and Earth Radiation Budget Satellite (ERBS) spacecraft. It is also shown that a pair of 3-min bilateration ranging transponder system (BRTS) tracking passes every 4 hrs, one each from two different BRTS locations, is sufficient to maintain user orbit accuracy to the navigation certification requirements. The BRTS tracking requirements for the single-TDRS configuration will also apply to each TDRS in a multiple-TDRS configuration.
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.
Nonorthogonal molecular orbital method: Single-determinant theory
NASA Astrophysics Data System (ADS)
Watanabe, Yoshihiro; Matsuoka, Osamu
2014-05-01
Using the variational principle, we have derived a variant of the Adams-Gilbert equation for nonorthogonal orbitals of a single-determinant wave function, which we name the modified Adams-Gilbert equation. If we divide the molecular system into several subsystems, such as bonds, lone pairs, and residues, we can solve the equations for the subsystems one by one. Thus, this procedure has linear scaling. We have presented a practical procedure for solving the equations that is also applicable to macromolecular calculations. The numerical examples show that the procedure yields, with reasonable effort, results comparable with those of the Hartree-Fock-Roothaan method for orthogonal orbitals. To resolve the convergence difficulty in the self-consistent-field iterations, we have found that virtual molecular-orbital shifts are very effective.
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.
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.
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.
Determination of Halo Orbits in the Sun-Jupiter System
NASA Astrophysics Data System (ADS)
Felipe, G.; Prado, A. F.
This paper has the goal of studying techniques to obtain Halo orbits in the Sun-Jupiter system. Halo orbits are special three-dimensional trajectories that exist around the Lagrangian points of the restricted three-body problem. These orbits are studied in several papers, since they have important applications in astronautics. To do that, an analytic calculation is performed using the Linstedt-Poincaré method. The basic idea of this method is to use the non-linear part of the differential equations written as a Legendre polynomial. After that, we know that when the non-linear terms of the differential equations are included in the variational equations, we have that the general solution will depend on the linear solution in the plane and in the z axis. Remember that in this case the motions are no longer separable. This method is then applied to the determination of the Halo orbits in the Sun-Jupiter system. Similar determinations has been made for the Sun-Earth and Earth-Moon systems in the literature.
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.
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.
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.
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.
Application of M-methods to satellite orbit determination
Xiaogang Pan; Haiyin Zhou; Yuanyuan Jiao; Weinan Jiang
2007-01-01
This paper analyses concretely the theory of M-methods, establishes the model of steady Satellite Orbit Determination (SOD) based on M-methods, which can farthest restrain gross errors so as to enhance the precision of SOD , and designs effective iterative algorithm named ameliorative Gauss-Newton method, by researching on the application condition of M-methods and the practical matter encountered in SOD. Finally
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.
Orbit determination for geostationary spacecraft using ion propulsion
Karlheinz Spindler
1994-01-01
The Advanced Relay and TEchnology MISsion (ARTEMIS) will be ESA's next geostationary communications-technology demonstration satellite. The inclination station keeping of this spacecraft is foreseen to be carried out by an ion propulsion system performing twice-daily low-thrust maneuvers, each with a duration of typically 3 hours. Therefore, orbit determination software had to be developed which can estimate, among other parameters, the
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.
Orbit determination with the tracking data relay satellite system
NASA Technical Reports Server (NTRS)
Argentiero, P.; Loveless, F.
1977-01-01
The possibility of employing the tracking data relay satellite system to satisfy the orbit determination demands of future applications missions is investigated. It is shown that when the relay satellites are continuously and independently tracked from ground stations it is possible, using six hour data arcs, to recover user satellite state with an average error of about 25 m radially, 260 m along track, and 20 m cross track. For this arc length, range sum data and range sum rate data are equally useful in determining orbits. For shorter arc lengths (20 min), range sum rate data is more useful than range sum data. When relay satellites are not continuously tracked, user satellite state can be recovered with an average error of about 140 m radially, 515 m along track, and 110 m cross track. These results indicate that the TDRS system can be employed to satisfy the orbit determination demands of applications missions, such as the MAGSAT and potential gradiometer missions, provided the relay satellites are continuously and independently tracked.
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.
Position determination systems. [using orbital antenna scan of celestial bodies
NASA Technical Reports Server (NTRS)
Shores, P. W. (inventor)
1976-01-01
A system for an orbital antenna, operated at a synchronous altitude, to scan an area of a celestial body is disclosed. The antenna means comprises modules which are operated by a steering signal in a repetitive function for providing a scanning beam over the area. The scanning covers the entire area in a pattern and the azimuth of the scanning beam is transmitted to a control station on the celestial body simultaneous with signals from an activated ground beacon on the celestial body. The azimuth of the control station relative to the antenna is known and the location of the ground beacon is readily determined from the azimuth determinations.
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.
Astrometric positioning and orbit determination of geostationary satellites
NASA Astrophysics Data System (ADS)
Montojo, F. J.; López Moratalla, T.; Abad, C.
2011-03-01
In the project titled “Astrometric Positioning of Geostationary Satellite” (PASAGE), carried out by the Real Instituto y Observatorio de la Armada (ROA), optical observation techniques were developed to allow satellites to be located in the geostationary ring with angular accuracies of up to a few tenths of an arcsec. These techniques do not necessarily require the use of large telescopes or especially dark areas, and furthermore, because optical observation is a passive method, they could be directly applicable to the detection and monitoring of passive objects such as space debris in the geostationary ring.By using single-station angular observations, geostationary satellite orbits with positional uncertainties below 350 m (2 sigma) were reconstructed using the Orbit Determination Tool Kit software, by Analytical Graphics, Inc. This software is used in collaboration with the Spanish Instituto Nacional de Técnica Aeroespacial.Orbit determination can be improved by taking into consideration the data from other stations, such as angular observations alone or together with ranging measurements to the satellite. Tests were carried out combining angular observations with the ranging measurements obtained from the Two-Way Satellite Time and Frequency Transfer technique that is used by ROA’s Time Section to carry out time transfer with other laboratories. Results show a reduction of the 2 sigma uncertainty to less than 100 m.
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.; So?nica, K.; Mervart, L.; Jäggi, A.
2015-05-01
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 z (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.
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.; So?nica, K.; Mervart, L.; Jäggi, 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.
Modeling radiation forces acting on satellites for precision orbit determination
NASA Technical Reports Server (NTRS)
Marshall, J. A.; Antreasian, P. G.; Rosborough, G. W.; Putney, B. H.
1992-01-01
Models of the TOPEX/Poseidon spacecraft are developed by means of finite-element analyses for use in generating acceleration histories for various orbit orientations which account for nonconservative radiation forces. The acceleration profiles are developed with an analysis based on the use of the 'box-wing' model in which the satellite is modeled as a combination of flat plates. The models account for the effects of solar, earth-albedo, earth-IR, and spacecraft-thermal radiation. The finite-element analysis gives the total force and induced accelerations acting on the satellite. The plate types used in the analysis have parameters that can be adjusted to optimize model performance according to the micromodel analysis and tracking observations. Acceleration related to solar radiation pressure is modeled effectively, and the techniques are shown to be useful for the precise orbit determinations required for spacecraft such as the TOPEX/Poseidon.
HOW TO DETERMINE AN EXOMOON'S SENSE OF ORBITAL MOTION
Heller, René [Origins Institute, McMaster University, Hamilton, ON L8S 4M1 (Canada); Albrecht, Simon, E-mail: rheller@physics.mcmaster.ca, E-mail: albrecht@phys.au.dk [Stellar Astrophysics Centre, Department of Physics and Astronomy, Aarhus University, Ny Munkegade 120, DK-8000 Aarhus C (Denmark)
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 ? Pic b causes an RME amplitude of almost 100 m s{sup –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.
Orbit/Attitude Determination and Control for the UMRSAT Mission
NASA Astrophysics Data System (ADS)
Dancer, M. W.; Searcy, J. D.
2008-08-01
As satellite missions become increasingly complex, a need for accurate determination and control systems using low cost hardware arises. This is especially true for university satellite programs such as the University of Missouri - Rolla satellite design team, or UMR SAT. With limited resources, mission success relies on creative and innovative hardware and software designs. This paper describes the development of control algorithms that will be used onboard the UMR SAT satellite pair. Using novel attitude and orbit control techniques and magnetometer-only attitude determination, the mission can be accomplished with low cost COTS hardware. The UMR developed ?-D controller will be used to facilitate the attitude and formation control, and the ?-D filter will be used for orbit determination. The ?-D technique has been successfully applied to a wide variety of applications ranging from wing aeroelastic flutter suppression to hit-to-kill missile autopilot design to reusable launch vehicle control. The results of each application have been very promising and show the potential improvement over pre-existing control techniques offered by the ?-D method. Along with software development, this paper also provides high fidelity simulations of the determination and control system are presented to demonstrate the effectiveness of the algorithms.
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2010 CFR
2010-04-01
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20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2014 CFR
2014-04-01
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20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2011 CFR
2011-04-01
...2011-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203...General § 416.203 Initial determinations of SSI eligibility. (a) What happens when you apply for SSI benefits. When you apply for...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2012 CFR
2012-04-01
...2012-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203...General § 416.203 Initial determinations of SSI eligibility. (a) What happens when you apply for SSI benefits. When you apply for...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2013 CFR
2013-04-01
...2013-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203...General § 416.203 Initial determinations of SSI eligibility. (a) What happens when you apply for SSI benefits. When you apply for...
Application of variance component estimation to precise orbit determination for ERS-2
NASA Astrophysics Data System (ADS)
Zhang, Fei-peng; Huang, Cheng; Feng, Chu-gang; Dong, Xiao-jun; Liao, Xin-hao
2001-10-01
Variance component estimation (VCE) is applied to precise orbit determination (POD) of the ERS-2 satellite. Twenty 5-day long arcs in the early three months in 1998 were calculated using the SLR and PRARE data. In the data the adjacent arcs overlap for two days except the intervals for orbit maneuver. The effect of VCE orbit determination on the calculation is investigated by an analysis of residuals and comparison of overlapping arcs, and the mean a posteriori standard deviation of each group of measured residuals is given. It is shown by the residuals analysis that the fitting of the measurements is significantly improved by VCD. However, according to Abbey criterion, VCD is not able to eliminate the systematic errors due to errors in the dynamic and geometric models. The results of the comparison of the overlapping arcs show that (1) VCE reduces the mean range deviation of overlapping arcs, especially where there are obviously unreasonable deviations, so that the orbit obtained has a more uniform precision; (2) By using VCE, adjacent arcs tend to close up and this is more apparent in the transverse direction. From the mean a posteriori standard error of each group of measurements, it can be seen that as far as the single normal point measurement is concerned, the data of some SLR stations are more important than other measurements in POD calculation. Generally speaking, determination of weighting by using VCE is more reasonable than by using initial standard deviation.
R. H. Gooding
1996-01-01
A new method has been developed for obtaining multiple solutions of the classical angles-only initial-orbit-determination problem. The method operates by a higher-order Newton correction of the assumed values for two of the unknown ranges, with the author's universal Lambert algorithm at the heart of the procedure. The observations are permitted to span several revolutions when the orbit is elliptic, and
R. H. Gooding
1996-01-01
A new method has been developed for obtaining multiple solutions of the classical angles-only initial-orbit-determination problem. The method operates by a higher-order Newton correction of the assumed values for two of the unknown ranges, with the author's universal Lambert algorithm at the heart of the procedure. The observations are permitted to span several revolutions when the orbit is elliptic, and
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.
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.
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.
Innovative observing strategy and orbit determination for Low Earth Orbit Space Debris
Milani, Andrea; Dimare, Linda; Rossi, Alessandro; Bernardi, Fabrizio
2011-01-01
We present the results of a large scale simulation, reproducing the behavior of a data center for the build-up and maintenance of a complete catalog of space debris in the upper part of the low Earth orbits region (LEO). The purpose is to determine the performances of a network of advanced optical sensors, through the use of the newest orbit determination algorithms developed by the Department of Mathematics of Pisa (DM). Such a network has been proposed to ESA in the Space Situational Awareness (SSA) framework by Carlo Gavazzi Space SpA (CGS), Istituto Nazionale di Astrofisica (INAF), DM, and Istituto di Scienza e Tecnologie dell'Informazione (ISTI-CNR). The conclusion is that it is possible to use a network of optical sensors to build up a catalog containing more than 98% of the objects with perigee height between 1100 and 2000 km, which would be observable by a reference radar system selected as comparison. It is also possible to maintain such a catalog within the accuracy requirements motivated by collisi...
Gregory B. Cook
1994-04-25
The construction of initial-data sets representing binary black-hole configurations in quasi-circular orbits is studied in the context of the conformal-imaging formalism. An effective-potential approach for locating quasi-circular orbits is outlined for the general case of two holes of arbitrary size and with arbitrary spins. Such orbits are explicitly determined for the case of two equal-sized nonrotating holes, and the innermost stable quasi-circular orbit is located. The characteristics of this innermost orbit are compared to previous estimates for it, and the entire sequence of quasi-circular orbits is compared to results from the post-Newtonian approximation. Some aspects of the numerical evolution of such data sets are explored.
Improved Space Object Orbit Determination Using CMOS Detectors
NASA Astrophysics Data System (ADS)
Schildknecht, T.; Peltonen, J.; Sännti, 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.
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.
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.
Preliminary orbit-determination method having no co-planar singularity
Robert M. L. Baker; Neil H. Jacoby
1977-01-01
It has long been recognized and demonstrated in the astrodynamic literature that three observations of angular position are not always sufficient to determine a preliminary orbit. One reason for this is due to the fact that as the plane of the observer's motion approaches the plane of the orbit of the observed object, the determination of the orbit of the
NASA Astrophysics Data System (ADS)
Jopek, T. J.; Rudawska, R.; Dybczynski, P. A.
2005-08-01
The value of the initial velocity of the stream meteoroids from the parent bodies is given by the physics of the outgassing of the cometary nuclei and by modeling the collisions between asteroids. In both cases the outflow speed of the meteoroid particles are small (Whipple 1951, Hughes 1977, 2000, Gustafson 1989, Jones 1995, Ma et al. 2002) and as result, the most meteoroid streams have similar orbits to either comets or asteroids. The formulae relating the changes of the orbital elements due to the small increment of the velocity were developed, among others by Plavec (1955), Pecina and Simek(1997), Williams (1996, 2001), Ma et al. (2001), Ma and Williams (2002). Assuming that the members of the observed meteor stream evolved dynamically under the influence of gravitational perturbations only, Pittich (1988), Harris and Hughes (1995), Williams (1996, 2001) estimated the initial velocity of the stream meteoroids. In their approach, Harris and Hughes have used the dispersion of the semimajor axes of the stream meteoroids. Williams proposed the method were used the mean orbit of the stream and the orbit of the identified parent body of the stream. The obtained results are not free from the discrepancy, explained partly by the particular orbital structure of the stream. However Kresak (1992) has strongly criticized the attempts to determine the initial velocities of the stream using the statistics of the meteor orbits. He argued that this is essentially impossible, because the dispersion of the initial velocities are masked by much larger measuring errors and also by the accumulated effects of planetary perturbations. In our paper, we decided to verify the reliability of the methods proposed by Harris and Hughes (1995), and by Williams (1996,2001). We made an numerical experiment consisting of the simulation of formation of several meteor streams and their dynamical evolution over 5000 years. We ejected meteoroids particles from the comets: Halley, Swift-Tuttle, Tempel-Tuttle. During the integration, the initial velocities of the stream members were estimated using the methods proposed by Hughes and Harris, and that by Williams. The results which we calculated till present, shows that the velocities obtained by the Williams method are to high when compared with the known velocities of the stream formation. On the other hand, the velocities obtained using Harris and Hughes method are to small. This work was supported by the KBN Project 2-P03-D-007-22.
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.
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.
Numerical black hole initial data with low eccentricity based on post-Newtonian orbital parameters
Walther, Benny; Bruegmann, Bernd; Mueller, Doreen
2009-06-15
Black hole binaries on noneccentric orbits form an important subclass of gravitational wave sources, but it is a nontrivial issue to construct numerical initial data with minimal initial eccentricity for numerical simulations. We compute post-Newtonian orbital parameters for quasispherical orbits using the method of Buonanno, Chen and Damour, (2006) and examine the resulting eccentricity in numerical simulations. Four different methods are studied resulting from the choice of Taylor-expanded or effective-one-body Hamiltonians, and from two choices for the energy flux. For equal-mass, nonspinning binaries the approach succeeds in obtaining low-eccentricity numerical initial data with an eccentricity of about e=0.002 for rather small initial separations of D > or approx. 10M. The eccentricity increases for unequal masses and for spinning black holes, but remains smaller than that obtained from previous post-Newtonian approaches. The effective-one-body Hamiltonian offers advantages for decreasing initial separation as expected, but in the context of this study also performs significantly better than the Taylor-expanded Hamiltonian for binaries with spin. For mass ratio 4 ratio 1 and vanishing spin, the eccentricity reaches e=0.004. For mass ratio 1 ratio 1 and aligned spins of size 0.85M{sup 2} the eccentricity is about e=0.07 for the Taylor method and e=0.014 for the effective-one-body method.
Miller's instability, microchaos and the short-term evolution of initially nearby orbits
Amina Helmi; Facundo Gomez
2007-10-02
We study the phase-space behaviour of nearby trajectories in integrable potentials. We show that the separation of nearby orbits initially diverges very fast, mimicking a nearly exponential behaviour, while at late times it grows linearly. This initial exponential phase, known as Miller's instability, is commonly found in N-body simulations, and has been attributed to short-term (microscopic) N-body chaos. However we show here analytically that the initial divergence is simply due to the shape of an orbit in phase-space. This result confirms previous suspicions that this transient phenomenon is not related to an instability in the sense of non-integrable behaviour in the dynamics of N-body systems.
Application of M-methods to satellite orbit determination
NASA Astrophysics Data System (ADS)
Pan, Xiaogang; Zhou, Haiyin; Jiao, Yuanyuan; Jiang, Weinan
2007-11-01
This paper analyses concretely the theory of M-methods, establishes the model of steady Satellite Orbit Determination (SOD) based on M-methods, which can farthest restrain gross errors so as to enhance the precision of SOD , and designs effective iterative algorithm named ameliorative Gauss-Newton method, by researching on the application condition of M-methods and the practical matter encountered in SOD. Finally the paper carries out simulation experiment to prove the validity of strategy put forward above. The results of experiment show that the algorithm can strongly improve the precision of the satellite, and markedly reduce the disadvantageous effect of gross errors. The algorithm can be realized easily in application to engineering.
Does a billiard orbit determine its (polygonal) table?
Bobok, Jozef
2008-01-01
We introduce a new equivalence relation on the set of all polygonal billiards. We say that two billiards (or polygons) are order equivalent if each of the billiards has an orbit whose footpoints are dense in the boundary and the two sequences of footpoints of these orbits have the same combinatorial order. We study this equivalence relation with additional regularity conditions on the orbit.
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.
Application of variance component estimation to precision orbit determination for ERS-2
NASA Astrophysics Data System (ADS)
Zhang, F. P.; Huang, C.; Feng, C. G.; Dong, X. J.; Liao, X. H.
2001-05-01
The variance component estimation (VCE) method was applied to the precision orbit determination (POD) for ERS-2 satellite. Totally, 23 5-day-long maneuver-free arcs with 2 days overlap (when without orbit maneuver separation between the successive arcs) in the early three months of 1998 were computed using SLR and PRARE data. To inspect the effect of the VCE method on the POD computation, the measurement residuals were analyzed, and the overlapping arcs were compared. At last, the mean posteriori standard deviations of each group of measurement were given. The analyses of residuals show that using the VCE method improved the fitness of measurements. Using Abbey criterion to examine systematic error checked the residuals. The checking results show that the VCE method could not eliminate the systematic errors in the orbit caused by the errors in dynamic and geometric models. The comparison results of overlapping arcs show that (1) using the VCE method reduced the mean deviations of overlapping arcs and the obviously unreasonable deviations in some arcs, thus made the orbit have more uniform precision; (2) relatively?the reduction of the mean deviations is more apparent in the transverse direction. From the mean posteriori standard deviations of each group of measurements, we could see that, for single normal points, the data of several SLR stations are more important than other measurements in POD computation. In general, the yielded mean posteriori standard deviations are more reasonable than the initial standard deviations to weight the observations.
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 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.
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.
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
Point-to-point sub-orbital space tourism: Some initial considerations
NASA Astrophysics Data System (ADS)
Webber, Derek
2010-06-01
Several public statements have been made about the possible, or even likely, extension of initial sub-orbital space tourism operations to encompass point-to-point travel. It is the purpose of this paper to explore some of the basic considerations for such a plan, in order to understand both its merits and its problems. The paper will discuss a range of perspectives, from basic physics to market segmentation, from ground segment logistics to spacecraft design considerations. It is important that these initial considerations are grasped before more detailed planning and design takes place.
TDRSS-user orbit determination using batch least-squares and sequential methods
D. H. Oza; T. L. Jones; M. Hakimi; Mina V. Samii; C. E. Doll; G. D. Mistretta; R. C. Hart
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
Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods
D. H. Oza; T. L. Jones; S. M. Fabien; G. D. Mistretta; R. C. Hart; C. E. Doll
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
Evaluation of Landsat4 orbit determination accuracy using batch least-squares and sequential methods
D. H. Oza; T. L. Jones; R. Feiertag; M. V. Samii; C. E. Doll; G. D. Mistretta; R. C. Hart
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
Improved solution accuracy for TDRSS-based TOPEX\\/Poseidon orbit determination
C. E. Doll; G. D. Mistretta; R. C. Hart; D. H. Oza; D. T. Bolvin; C. M. Cox; M. Nemesure; D. J. Niklewski; M. V. Samii
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
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.
A fuzzy clustering application to precise orbit determination
NASA Astrophysics Data System (ADS)
Soto, Jesus; Vigo Aguiar, M. Isabel; Flores-Sintas, Antonio
2007-07-01
In recent years, fuzzy logic techniques have been successfully applied in geodesy problems, in particular to GPS. The aim of this work is to test a fuzzy-logic method with an enhanced probability function as a tool to provide a reliable criteria for weighting scheme for satellite-laser-ranging (SLR) station observations, seeking to optimize their contribution to the precise orbit determination (POD) problem. The data regarding the stations were provided by the International Laser Ranging Service (ILRS), NASA/Crustal Dynamics Data Information System (CDDIS) provided the satellite data for testing the method. The software for processing the data is GEODYN II provided by NASA/Goddard Space Flight Center (GSFC). Factors to be considered in the fuzzy-logic clustering are: the total number of LAGEOS passes during the past 12 months, the stability measure of short- and long-term biases, the percentage of LAGEOS normal points that were accepted in CSR weekly LAGEOS analysis, and the RMS uncertainty of the station coordinates. A fuzzy-logic statistical method allows classifying the stations through a clear `degree of belonging' to each station group. This degree of belonging translates into a suitable weight to be assigned to each station in the global solutionE The first tests carried out showed improvements in the RMS of the global POD solution as well as individual stations, to within a few millimeters. We expect further work would lead to further improvements.
Determining the eccentricity of the Moon's orbit without a telescope
NASA Astrophysics Data System (ADS)
Krisciunas, Kevin
2010-08-01
Prior to the invention of the telescope many astronomers worked out models of the motion of the Moon to predict the position of the Moon in the sky. These geometrical models implied a certain range of distances of the Moon from Earth. Ptolemy's most quoted model predicted that the Moon was nearly twice as far away at apogee than at perigee. Measurements of the angular size of the Moon were within the capabilities of pretelescopic astronomers. Such measurements could have helped refine the models of the motion of the Moon, but hardly anyone seems to have made any measurements that have come down to us. We use a piece of cardboard with a small hole in it which slides up and down a yardstick to show that it is possible to determine the eccentricity ?~0.039+/-0.006 of the Moon's orbit. A typical measurement uncertainty of the Moon's angular size is +/-0.8 arc min. Because the Moon's angular size ranges from 29.4 to 33.5 arc min, carefully taken naked eye data are accurate enough to demonstrate periodic variations of the Moon's angular size.
Low-cost autonomous orbit control about Mars: Initial simulation results
NASA Astrophysics Data System (ADS)
Dawson, S. D.; Early, L. W.; Potterveld, C. W.; Königsmann, H. J.
1999-11-01
Interest in studying the possibility of extraterrestrial life has led to the re-emergence of the Red Planet as a major target of planetary exploration. Currently proposed missions in the post-2000 period are routinely calling for rendezvous with ascent craft, long-term orbiting of, and sample-return from Mars. Such missions would benefit greatly from autonomous orbit control as a means to reduce operations costs and enable contact with Mars ground stations out of view of the Earth. This paper present results from initial simulations of autonomously controlled orbits around Mars, and points out possible uses of the technology and areas of routine Mars operations where such cost-conscious and robust autonomy could prove most effective. These simulations have validated the approach and control philosophies used in the development of this autonomous orbit controller. Future work will refine the controller, accounting for systematic and random errors in the navigation of the spacecraft from the sensor suite, and will produce prototype flight code for inclusion on future missions. A modified version of Microcosm's commercially available High Precision Orbit Propagator (HPOP) was used in the preparation of these results due to its high accuracy and speed of operation. Control laws were developed to allow an autonomously controlled spacecraft to continuously control to a pre-defined orbit about Mars with near-optimal propellant usage. The control laws were implemented as an adjunct to HPOP. The GSFC-produced 50 × 50 field model of the Martian gravitational potential was used in all simulations. The Martian atmospheric drag was modeled using an exponentially decaying atmosphere based on data from the Mars-GRAM NASA Ames model. It is hoped that the simple atmosphere model that was implemented can be significantly improved in the future so as to approach the fidelity of the Mars-GRAM model in its predictions of atmospheric density at orbital altitudes. Such additional work would take the form of solar flux (F10.7) and diurnal density dependencies. The autonomous controller is a-derivative of the proprietary and patented Microcosm Earth-orbiting control methodology which will be implemented on the upcoming Surrey Satellite Technology (SSTL) UoSAT-12 and the NASA EO-1 spacecraft missions. This work was funded by the NASA Jet Propulsion Laboratory under a Phase I SBIR (96.1 07.02 9444) and by internal Microcosm R&D funds as well as earlier supporting work done under a variety of USAF Research Laboratory-sponsored contracts [1, 2, 4, 12].
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.
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.
Rui Guo; Xiaogong Hu; Li Liu; Xiaoli Wu; Yong Huang; Feng He
2010-01-01
Geostationary satellites (GEOs) play a significant role in the regional satellite navigation system. Simulation experiments\\u000a show that the clock corrections could be mitigated through a single strategy or double differencing strategies for a navigation\\u000a constellation, but for the mode of individual GEO orbit determination, high precision orbit and clock correction could not\\u000a be obtained in the orbit determination based on
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 ±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 ±0.8 arcmin can be achieved.
Advanced stellar compass onboard autonomous orbit determination, preliminary performance.
Betto, Maurizio; Jørgensen, John L; Jørgensen, 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
Orbital maneuvers between halo orbits
NASA Astrophysics Data System (ADS)
Prado, A.; Felipe, G.
This paper has the goal of studying the problem of orbital transfers between two Halo orbits Halo orbits are special three-dimensional trajectories that exist around the Lagrangian points of the restricted three-body problem These orbits are studied in several papers since they have important applications in astronautics The first step involved in this research is to perform the determination of the Halo orbits To do that an analytic calculation is performed using the Linstedt-Poincar e method The present paper considers that a maneuver will be performed to transfer the spacecraft from an initial to a final Halo orbit The control that will be used to achieve that goal is constituted by a pair of instantaneous change in the velocity of the spacecraft at the beginning and at the end of the maneuver A numerical algorithm based in the Lambert Problem is built to calculate the transfer orbits The two orbits are divided in several points and the algorithm is applied to each pair of points Finally the solution that gives the minimum fuel consumption is found This maneuver can be used to i change the orbit of the spacecraft to allow a second application in a different Halo orbit ii to perform station keeping in a Halo orbit that is escaping from its nominal orbital parameters
Precise orbit determination of GLONASS satellites at the European space agency
I. Romero; C. Garcia; R. Kahle; J. Dow; T. Martin-Mur
2002-01-01
As an active Analysis Centre of the International GPS Service (IGS) the European Space Operations Centre (ESOC) joined the IGEX program for precise orbit determination of the GLONASS satellite constellation since its inception in 1998. This paper describes the orbit determination processing strategy, the specific GLONASS modelling issues implemented and a discussion of the processing results.
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.
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2014 CFR
2014-10-01
...instructions. When a paper RA is mailed, it must comply with CMS manual instructions that parallel the HIPAA data content and coding requirements. (2) The notice of initial determination must contain: (i) The basis for any full or partial denial...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2011 CFR
2011-10-01
...instructions. When a paper RA is mailed, it must comply with CMS manual instructions that parallel the HIPAA data content and coding requirements. (2) The notice of initial determination must contain: (i) The basis for any full or partial denial...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2012 CFR
2012-10-01
...instructions. When a paper RA is mailed, it must comply with CMS manual instructions that parallel the HIPAA data content and coding requirements. (2) The notice of initial determination must contain: (i) The basis for any full or partial denial...
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2013 CFR
2013-10-01
...instructions. When a paper RA is mailed, it must comply with CMS manual instructions that parallel the HIPAA data content and coding requirements. (2) The notice of initial determination must contain: (i) The basis for any full or partial denial...
20 CFR 410.620 - Notice of initial determination.
Code of Federal Regulations, 2011 CFR
2011-04-01
...FEDERAL COAL MINE HEALTH AND SAFETY ACT OF 1969, TITLE IV-BLACK LUNG BENEFITS (1969- ) Determinations of Disability...party's entitlement to benefits has ended because of such party's death (see § 410.610(c)). If the initial...
20 CFR 410.621 - Effect of initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 410.621 Section 410.621 Employees' Benefits SOCIAL SECURITY ADMINISTRATION FEDERAL COAL MINE HEALTH AND SAFETY...Determinations, Administrative Review, Finality of Decisions, and Representation of Parties § 410.621 Effect of initial...
Spiga, Aymeric
Initial results from radio occultation measurements with the Mars Reconnaissance Orbiter, dynamics Meteorology a b s t r a c t The Mars Reconnaissance Orbiter (MRO) performs radio occultation (RO (MRO) has been performing radio occultation (RO) measurements since the primary science phase
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.
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.
NASA Astrophysics Data System (ADS)
Doll, C.; Mistretta, G.; Hart, R.; Oza, D.; Cox, C.; Nemesure, M.; Bolvin, D.; Samii, Mina V.
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.
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.
Orbit Determination of the EnVision D-InSAR mission to Venus
NASA Astrophysics Data System (ADS)
Cochrane, C. G.; Ghail, R.; Wilson, C. F.; Hall, D.; Mason, P.; EnVision Mission Design Team
2011-12-01
EnVision is proposed to the European Space Agency for launch in 2020-22, to detect geological change on Venus and to address key questions from Magellan and Venus Express, particularly: how active is the surface of Venus now? The global distribution of impact craters has been shown (Romeo and Turcotte, 2010) as consistent with a range of resurfacing models, with rates of volcanic and tectonic activity comparable to that of plate interiors on Earth, i.e., perhaps up to 10 mm a-1 in localised tectonic movement and less than 1km3 a-1 (Stofan et al., 2005) in global extrusive activity. Capturing, during the lifetime of a single spacecraft, evidence of an individual volcanic eruption or tectonic movement under the clouds of Venus is only feasible using Differential Interferometric SAR (D-InSAR). D-InSAR is a radar interferometry technique using repeat-pass images of a surface area to measure terrain deformation using radar phase differences at the beginning and end of the observation period. To separate the effects of topography, at least three images are required: a pair closely spaced in time, in which any phase differences are solely attributable to topography, and a third acquired later but from one of the orbital positions used previously. Such technology is now routinely used in earth orbit for terrestrial applications. EnVision is a mission that will use these techniques at Venus to determine the rate and location of geological activity (e.g., tectonics, earthquakes, or volcanic eruptions) on Venus. The EnVision radar will be designed to detect rates of movement as close to the theoretical limit of 1 mm a-1as possible, from a quasi-circular, quasi-polar orbit of 300 km altitude, over a period of up to 10 years. Hence, EnVision must repeat its initial orbit throughout its 10 year mission and in synchronism with surface rotation to within 100 m in all three axes. This paper describes how we propose to determine EnVision's orbit to the necessary degree of accuracy by cost-effective use of the techniques listed in the table.
C. Doll; G. Mistretta; R. Hart; D. Oza; C. Cox; M. Nemesure; D. Bolvin; Mina V. Samii
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
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.
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.
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.
NASA Astrophysics Data System (ADS)
Oza, D. H.; Jones, T. L.; Feiertag, R.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
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.
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.
Analysis of orbit determination for space based optical space surveillance system
NASA Astrophysics Data System (ADS)
Sciré, 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.
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.
Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods
NASA Astrophysics Data System (ADS)
Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.
1991-10-01
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.
Orbit Determination with the two-body Integrals. III
Giovanni F. Gronchi; Giulio Bau'; Stefano Maro'
2015-01-15
We present the results of our investigation on the use of the two-body integrals to compute preliminary orbits by linking too short arcs of observations of celestial bodies. This work introduces a significant improvement with respect to the previous papers on the same subject (see Gronchi et al. 2010, 2011). Here we find a univariate polynomial equation of degree 9 in the radial distance $\\rho$ of the orbit at the mean epoch of one of the two arcs. This is obtained by a combination of the algebraic integrals of the two-body problem. Moreover, the elimination step, which in Gronchi et al. 2010, 2011 was done by resultant theory coupled with the discrete Fourier transform, is here obtained by elementary calculations. We also show some numerical tests to illustrate the performance of the new algorithm.
GLONASS Orbit Determination at the Center for Space Research
R. J. Eanes; R. S. Nerem; P. A. M Abusali; W. Bamford; J. C. Ries
In this article we review the contributions of the Center for Space Research to the International GLONASS Experiment 98 (IGEX-98) campaign. These include 1) the temporary establishment of a GPS\\/GLONASS receiver at the IGS GPS site at McDonald Observatory in west Texas; 2) the evaluation of GLONASS orbits computed by different centers using Satellite Laser Ranging (SLR) data; 3) the
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.
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.
Determinant Quantum Monte Carlo Study of the Orbitally Selective Mott Transition K. Bouadim,1
Scalettar, Richard T.
Determinant Quantum Monte Carlo Study of the Orbitally Selective Mott Transition K. Bouadim,1 G. G that an orbitally selective Mott transition (OSMT) occurs in which the more weakly interacting band can be metallic is at the opposite extreme: it has no hopping from site to site (zero bandwidth) and instead has only an on-site
The orbit determination of (4179) Toutatis from optical and radar data
N. V. Krivova; E. I. Yagudina; V. A. Shor
1994-01-01
The conventional technique of the least squares fit was used for the orbit determination of (4179) Toutatis on the basis of all available data: optical observations from February 1934 to February 1993 and radar data (delay and Doppler), obtained in 1992 during the asteroid approach to the Earth. The improved orbit and uncertainties analysis are presented.
From Astrometry to Celestial Mechanics: Orbit Determination with Very Short Arcs
Andrea Milani; Zoran Knezevic
2005-01-01
Contemporary surveys provide a huge number of detections of small solar system bodies, mostly asteroids. Typically, the reported astrometry is not enough to compute an orbit and\\/or perform an identification with an already discovered object. The classical methods for preliminary orbit determination fail in such cases: a new approach is necessary. When the observations are not enough to compute an
Derivation of a general perturbation solution - Its application to determination of orbit
NASA Technical Reports Server (NTRS)
Born, G. H.
1970-01-01
Analytical solution to three-body problems is applied to the problem of predicting the orbit of a lunar satellite and determining the orbit of a near-earth satellite. Using this solution, the state vector may be generated at any time without intermediate numerical extrapolation.
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.
Improved solution accuracy for TDRSS-based TOPEX/Poseidon orbit determination
NASA Astrophysics Data System (ADS)
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.
1994-05-01
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.
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.
Orbit Determination with Topocentric Correction: Algorithms for the Next Generation Surveys
Andrea Milani; Giovanni F. Gronchi; Davide Farnocchia; Zoran Knezevic; Robert Jedicke; Larry Denneau; Francesco Pierfederici
2007-07-28
Given a set of astrometric observations of the same object, the problem of orbit determination is to compute the orbit and to assess its uncertainty and reliability. For the next generation surveys, with much larger number density of observed objects, new algorithms or substantial revisions of the classical ones are needed. The problem has three main steps, preliminary orbit, least squares orbit, and quality control. The classical theory of preliminary orbits was incomplete: the consequences of the topocentric correction had not been fully studied. We show that it is possible to account for this correction, possibly with an increase in the number of preliminary solutions, without impairing the overall orbit determination performance. We have developed modified least squares orbit determination algorithms that can be used to improve the reliability of the procedure. We have tested the complete procedure on two simulations with number densities comparable to that expected from the next generation surveys such as Pan-STARRS and LSST. To control the problem of false identifications we have introduced a quality control on the fit residuals based on an array of metrics and a procedure to remove duplications and contradictions in the output. The results confirm that large sets of discoveries can be obtained with good quality orbits and very high success rate losing only 0.6 to 1.3% of objects and a false identification rate in the range 0.02 to 0.06%.
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.
Hubble Space Telescope: Cosmic Origins Spectrograph FUV detector initial on-orbit performance
NASA Astrophysics Data System (ADS)
McPhate, Jason B.; Siegmund, Oswald H.; Vallerga, John V.; Sahnow, David J.; Ake, Thomas B.; Penton, Steven V.; France, Kevin; Massa, Derck; Osterman, Steven N.; Béland, Stéphane; McCandliss, Stephan R.
2010-07-01
The Cosmic Origins Spectrograph (COS) was installed on the Hubble Space Telescope (HST) in May 2009 during Servicing Mission 4 (SM4). This paper discusses the initial on-orbit performance of the HST-COS far ultraviolet (FUV) detector designed and built by the Experimental Astrophysics Group at the Univ. of California, Berkeley. The HST-COS FUV detector is an open face, photon counting, microchannel plate (MCP) based device employing a cross delay line (XDL) readout. The detector consists of two separate, end-to-end segments (2x 85mm x 10mm - 179mm x 10mm total with a gap between segments), each digitized within a 16384x1024 space. The input surface is curved to match the Rowland circle of HST-COS. The CsI photocathode and open face nature result in sensitivity from <900Å to ~1750Å. Spatial resolution is approximately 25-30?m. Comparisons of on-orbit behavior relative to expectations from ground testing are performed. Areas of discussion include background (rate and morphology), sensitivity (system throughput and short wavelength response), and imaging performance (apparent spatial resolution and flat field fixed pattern). A measured increase in the MCP gain relative to ground testing is also discussed.
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.
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.
Orbit on demand - Will cost determine best design?
NASA Technical Reports Server (NTRS)
Macconochie, J. O.; Mackley, E. A.; Morris, S. J.; Phillips, W. P.; Breiner, C. A.; Scotti, S. J.
1985-01-01
Eleven design concepts for vertical (V) and horizontal (H) take-off launch-on-demand manned orbital vehicles are discussed. Attention is given to up to three stages, Mach numbers (sub-, 2, or 3), expendable boosters, drop tanks (DT), and storable (S) or cryogenic fuels. All the concepts feature lifting bodies with circular cross-section and most have a 7 ft diam, 15 ft long payload bay as well as a crew compartment. Expendable elements impose higher costs and in some cases reduce all-azimuth launch capabilities. Single-stage vehicles simplify the logistics whether in H or V configuration. A two-stage H vehicle offers launch offset for the desired orbital plane before firing the rocket engines after take-off and subsonic acceleration. A two-stage fully reusable V form has the second lowest weight of the vehicles studied and an all-azimuth launch capability. Better definition of the prospective mission requirements is needed before choosing among the alternatives.
The Suomi National Polar-orbiting Partnership Mission: Status and Initial Results
NASA Astrophysics Data System (ADS)
Gleason, J. F.; Butler, J. J.; Hsu, N. C.
2012-12-01
The Suomi National Polar-orbiting Partnership ( S-NPP) was launched October 28, 2011. After an initial checkout and activation period, all the instruments are in nominal operation and data products are being publicly released. Suomi NPP has five instruments, VIsible-InfraRed Radiometer Suite(VIIRS), Cross-track Infrared Sounder(CrIS), Advanced Technology Microwave Sounder ATMS, Ozone Mapping and Profiler Suite (OMPS), and Clouds And The Earth's Radiant Energy System (CERES). Suomi NPP will provide 28 data products to two primary user communities; operational users in the weather and environmental conditions forecasting and scientific users primarily interested in studying longer-term climate processes. This presentation will focus on the current status of the Suomi NPP mission, instruments and data products.
The determinants of MNC subsidiary initiatives: implications for small business
Cher-Hung Tseng; Cher-Min Fong; Kuo-Hsien Su
2004-01-01
Since the importance of the subsidiary's role continues to increase, a growing number of studies have focused on MNC subsidiary strategies. The aim of this study is to explore the determinants of a subsidiary's initiative. Based on the subsidiary research's classification provided by Birkinshaw and Hood, an integrated framework is developed to examine the influences of three groups of variables
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.
Orbit Determination and Differential-drag Control of Planet Labs Cubesat Constellations
Foster, Cyrus; Mason, James
2015-01-01
We present methodology and mission results from orbit determination of Planet Labs nanosatellites and differential-drag control of their relative motion. Orbit determination (OD) is required on Planet Labs satellites to accurately predict the positioning of satellites during downlink passes and we present a scalable OD solution for large fleets of small satellites utilizing two-way ranging. In the second part of this paper, we present mission results from relative motion differential-drag control of a constellation of satellites deployed in the same orbit.
Study of geopotential error models used in orbit determination error analysis
NASA Technical Reports Server (NTRS)
Yee, C.; Kelbel, D.; Lee, T.; Samii, M. V.; Mistretta, G. D.; Hart, R. C.
1991-01-01
The uncertainty in the geopotential model is currently one of the major error sources in the orbit determination of low-altitude Earth-orbiting spacecraft. The results of an investigation of different geopotential error models and modeling approaches currently used for operational orbit error analysis support at the Goddard Space Flight Center (GSFC) are presented, with emphasis placed on sequential orbit error analysis using a Kalman filtering algorithm. Several geopotential models, known as the Goddard Earth Models (GEMs), were developed and used at GSFC for orbit determination. The errors in the geopotential models arise from the truncation errors that result from the omission of higher order terms (omission errors) and the errors in the spherical harmonic coefficients themselves (commission errors). At GSFC, two error modeling approaches were operationally used to analyze the effects of geopotential uncertainties on the accuracy of spacecraft orbit determination - the lumped error modeling and uncorrelated error modeling. The lumped error modeling approach computes the orbit determination errors on the basis of either the calibrated standard deviations of a geopotential model's coefficients or the weighted difference between two independently derived geopotential models. The uncorrelated error modeling approach treats the errors in the individual spherical harmonic components as uncorrelated error sources and computes the aggregate effect using a combination of individual coefficient effects. This study assesses the reasonableness of the two error modeling approaches in terms of global error distribution characteristics and orbit error analysis results. Specifically, this study presents the global distribution of geopotential acceleration errors for several gravity error models and assesses the orbit determination errors resulting from these error models for three types of spacecraft - the Gamma Ray Observatory, the Ocean Topography Experiment, and the Cosmic Background Explorer.
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.
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.
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.
Precise orbit determination of Compass-M1: a primary result
NASA Astrophysics Data System (ADS)
Sun, Baoqi
On April 13, 2007, the first experiment satellite, Compass-M1, of China's the second generation Compass Navigation system was successfully launched. Unlike previous Compass satellites, Compass-M1 is the first satellite in medium earth orbit (MEO), and broadcast navigation signals in multi-frequencies in L-band. If signals were received from more than four satellites, users can determine their locations in a passive manner like using GPS. A primary result of precise orbit determination of Compass-M1 is presented in this paper. Five tracking stations, all located in China, are used. Double-frequency code and carrier phase observations are processed in zero-difference mode. Receiver and satellite clocks are modeled by linear or quadratic polynomial. The radiation pressure model is the so-called extended CODE orbit model, and an a priori model is introduced according to the size and physical attribute of Compass-M1. The solution is based on 3-day arc dynamical precise orbit determination. Estimated parameters include six keplerian orbit elements, two radiation pressure model parameters and clock polynomial coefficients. Orbit overlap difference and validating with SLR indicate that the accuracy of the precise orbit is quite exciting and exceeds our expectation.
Determining final probabilities directly from the initial state
Sanz, A S
2013-01-01
In quantum scattering problems it is not possible to relate unambiguously a particular feature of the final outcome with some specific section of the initial state. As it is shown here, this drawback can be overcome by conveniently combining the divergence or Gauss-Ostrogradsky theorem with Bohmian mechanics. This renders a general approach which enables such a connection and allows us to determine the value of final partial or restricted probabilities directly from the corresponding localized section of the initial state. As an illustration, this approach is applied to two prototypical scattering phenomena: tunneling and grating diffraction.
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.
NASA Astrophysics Data System (ADS)
Escher, William J. D.
1992-08-01
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.
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.
G. M. Keating; S. W. Bougher; M. E. Theriot; R. H. Tolson; R. C. Blanchard; R. W. Zurek; J. M. Forbes; J. Murphy
2006-01-01
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
Onboard orbit determination using GPS observations based on the unscented Kalman filter
NASA Astrophysics Data System (ADS)
Choi, Eun-Jung; Yoon, Jae-Cheol; Lee, Byoung-Sun; Park, Sang-Young; Choi, Kyu-Hong
2010-12-01
Spaceborne GPS receivers are used for real-time navigation by most low Earth orbit (LEO) satellites. In general, the position and velocity accuracy of GPS navigation solutions without a dynamic filter are 25 m (1 ?) and 0.5 m/s (1 ?), respectively. However, GPS navigation solutions, which consist of position, velocity, and GPS receiver clock bias, have many abnormal excursions from the normal error range for space operation. These excursions lessen the accuracy of attitude control and onboard time synchronization. In this research, a new onboard orbit determination algorithm designed with the unscented Kalman filter (UKF) was developed to improve the performance. Because the UKF is able to obtain the posterior mean and covariance accurately by using the second-order Taylor series expansion through the sampled sigma points that are propagated by using the true nonlinear system, its performance can be better than that of the extended Kalman filter (EKF), which uses the linearized state transition matrix to predict the covariance. The dynamic models for orbit propagation applied perturbations due to the 40 × 40 geo-potential, the gravity of the Sun and Moon, solar radiation pressure, and atmospheric drag. The 7(8)th-order Runge-Kutta numerical integration was applied for orbit propagation. Two types of observations, navigation solutions and C/A code pseudorange, can be used at the user's discretion. The performances of the onboard orbit determination were verified using real GPS data of the CHAMP and KOMPSAT-2 satellites. The results of the orbit determination were compared with the precision orbit ephemeris (POE) of the CHAMP and KOMPSAT-2 satellites. The comparison of the orbit determination results using EKF and UKF shows that orbit determination using the UKF yields better results than that using the EKF. In addition, the estimation of the accuracy using the C/A code pseudorange is better than that using the navigation solutions. The absolute position and velocity accuracies of the UKF using GPS C/A code pseudorange were 12.098 m and 0.0159 m/s in the case of the CHAMP satellite, and 8.172 m and 0.0085 m/s in the case of the KOMPSAT-2 satellite. Moreover, the abnormal excursions of navigation solutions can be eliminated. These results verify that onboard orbit determination using GPS C/A code pseudorange, which is based on the UKF can provide more stable and accurate orbit information in the spaceborne GPS receiver.
Federal Register 2010, 2011, 2012, 2013, 2014
2012-07-06
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Long arc orbit determination for Mariner 9 using combined radio and optical data types
NASA Technical Reports Server (NTRS)
Born, G. H.; Mohan, S. N.
1974-01-01
Results of postflight orbit determination for the Mariner 9 trajectory (Mariner-Mars 1971 mission) using Doppler data combined with optical data in the form of TV photographs of the natural satellites and Mars' surface features. It is shown that the combined data yield a solution which is a factor of 3 to 5 better than that obtained from Doppler-only solutions for the node of the orbit relative to the plane perpendicular to the Earth-Mars line. Consequently, these optical data types have a demonstrated potential to significantly improve planetary orbiter navigation accuracies.
From Astrometry to Celestial Mechanics: Orbit Determination with Very Short Arcs
Andrea Milani; Zoran Kneževi?
Contemporary surveys provide a huge number of detections of small solar system bodies, mostly asteroids. Typically, the reported\\u000a astrometry is not enough to compute an orbit and\\/or perform an identification with an already discovered object. The classical\\u000a methods for preliminary orbit determination fail in such cases: a new approach is necessary. When the observations are not\\u000a enough to compute an
L. A. Campbell; J. W. Cook; K. E. Cunningham; W. A. Feess
1988-01-01
The authors consider the orbit-determination square-root information filter\\/smoother (SRIF\\/S) component for TRACE, a large orbital analysis program. They review the TRACE SRIF\\/S design rationale and implementation effort from the perspective of a half decade of experience with its practical use, concentrating on the features that differentiate the TRACE implementation from others. This presentation is a largely anecdotal account of what
Precise GLONASS orbit determination within the IGS\\/IGLOS – Pilot Project
Robert Weber; James A. Slater; Elisabeth Fragner; Vladimir Glotov; Heinz Habrich; Ignacio Romero; Stefan Schaer
2005-01-01
This paper reviews the reasons for setting up the IGS GLONASS Pilot Project and the objectives of this initiative. Then we present the operational tracking network as well as the current state of the GLONASS constellation. Afterwards, we discuss in more detail the generation and consistency of current GLONASS precise orbits based on microwave and on laser tracking data. Later
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.
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.
Orbit and attitude determination results during launch support operations for SBS-5
NASA Technical Reports Server (NTRS)
Hartman, K. R.; Iano, P. J.
1989-01-01
Presented are orbit and attitude determination results from the launch of Satellite Business Systems (SBS)-5 satellite on September 8, 1988 by Arianespace. SBS-5 is a (HS-376) spin stabilized spacecraft. The launch vehicle injected the spacecraft into a low inclination transfer orbit. Apogee motor firing (AMF) attitude was achieved with trim maneuvers. An apogee kick motor placed the spacecraft into drift orbit. Postburn, reorientation and spindown maneuvers were performed during the next 25 hours. The spacecraft was on-station 19 days later. The orbit and attitude were determined by both an extended Kalman filter and a weighted least squares batch processor. Although the orbit inclination was low and the launch was near equinox, post-AMF analysis indicated an attitude declination error of 0.034 deg., resulting in a saving of 8.5 pounds of fuel. The AMF velocity error was 0.4 percent below nominal. The post-AMF drift rate was determined with the filter only 2.5 hours after motor firing. The filter was used to monitor and retarget the reorientation to orbit normal in real time.
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.
A high order method for orbital conjunctions analysis: Sensitivity to initial uncertainties
NASA Astrophysics Data System (ADS)
Morselli, Alessandro; Armellin, Roberto; Di Lizia, Pierluigi; Bernelli Zazzera, Franco
2014-02-01
A high order method to quickly assess the effect that uncertainties produce on orbital conjunctions through a numerical high-fidelity propagator is presented. In particular, the dependency of time and distance of closest approach to initial uncertainties on position and velocity of both objects involved in a conjunction is studied. The approach relies on a numerical integration based on differential algebraic techniques and a high-order algorithm that expands the time and distance of closest approach in Taylor series with respect to relevant uncertainties. The modeled perturbations are atmospheric drag, using NRLMSISE-00 air density model, solar radiation pressure with shadow, third body perturbation using JPL's DE405 ephemeris, and EGM2008 gravity model. The polynomial approximation of the final position is used as an input to compute analytically the expansion of time and distance of closest approach. As a result, the analysis of a close encounter can be performed through fast, multiple evaluations of Taylor polynomials. Test cases with objects ranging from LEO to GEO regimes are considered to assess the performances and the accuracy of the proposed method.
NASA Astrophysics Data System (ADS)
Wu, Jinjie; Liu, Kun; Wei, Jingbo; Han, Dapeng; Xiang, Junhua
2012-12-01
Particle filter (PF) is widely used in nonlinear and non-Gaussian systems. Resampling is one of the significant steps in PF. However, PF using conventional resampling approaches may lead to divergent solutions because of the degeneracy phenomenon or sample impoverishment associated with a multidimensional system. In this article, an efficient alternative to conventional resampling approaches, called adaptive partial systematic resampling (APSR) with Markov chain Monte Carlo move and intelligent roughening is proposed for satellite orbit determination using a magnetometer. The results of the new resampling approach are compared with conventional resampling approaches and with unscented Kalman filter (UKF) for various initial errors in position and velocity, measurement sampling periods, and measurement noises to evaluate and verify the performance of the new resampling approach. The results of the new resampling approach in all cases are significantly better than the results of conventional resampling approaches. The velocity accuracy of the orbit determination of APSR is slightly poorer than UKF for relatively small initial errors, and small Gaussian measurement noise. However, the proposed approach yields more robust and stable convergence than UKF under large initial errors, long measurement sampling period, large Gaussian measurement noise, or non-Gaussian noise.
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.
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.
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.
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
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.
NASA Astrophysics Data System (ADS)
Herz, A.; Stoner, F.
2013-09-01
Current SSA sensor tasking and scheduling is not centrally coordinated or optimized for either orbit determination quality or efficient use of sensor resources. By applying readily available capabilities for determining optimal tasking times and centrally generating de-conflicted schedules for all available sensors, both the quality of determined orbits (and thus situational awareness) and the use of sensor resources may be measurably improved. This paper provides an approach that is logically separated into two main sections. Part 1 focuses on the science of orbit determination based on tracking data and the approaches to tracking that result in improved orbit prediction quality (such as separating limited tracking passes in inertial space as much as possible). This part of the paper defines the goals for Part 2 of the paper which focuses on the details of an improved tasking and scheduling approach for sensor tasking. Centralized tasking and scheduling of sensor tracking assignments eliminates conflicting tasking requests up front and coordinates tasking to achieve (as much as possible within the physics of the problem and limited resources) the tracking goals defined in Part I. The effectivity of the proposed approach will be assessed based on improvements in the overall accuracy of the space catalog. Systems Tool Kit (STK) from Analytical Graphics and STK Scheduler from Orbit Logic are used for computations and to generate schedules for the existing and improved approaches.
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.
42 CFR 422.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2013 CFR
2013-10-01
...2013-10-01 false Notice and effect of initial determinations...Appeal procedures for Civil Money Penalties § 422.1018 Notice and effect of initial determinations...for the determination, the effect of the determination,...
42 CFR 423.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2013 CFR
2013-10-01
...2013-10-01 false Notice and effect of initial determinations...Appeal Procedures for Civil Money Penalties § 423.1018 Notice and effect of initial determinations...for the determination, the effect of the determination,...
42 CFR 422.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2011 CFR
2011-10-01
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42 CFR 423.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2012 CFR
2012-10-01
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42 CFR 422.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2014 CFR
2014-10-01
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42 CFR 423.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2011 CFR
2011-10-01
...2011-10-01 false Notice and effect of initial determinations...Appeal Procedures for Civil Money Penalties § 423.1018 Notice and effect of initial determinations...for the determination, the effect of the determination,...
42 CFR 423.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2014 CFR
2014-10-01
...2014-10-01 false Notice and effect of initial determinations...Appeal Procedures for Civil Money Penalties § 423.1018 Notice and effect of initial determinations...for the determination, the effect of the determination,...
42 CFR 422.1018 - Notice and effect of initial determinations.
Code of Federal Regulations, 2012 CFR
2012-10-01
...2012-10-01 false Notice and effect of initial determinations...Appeal procedures for Civil Money Penalties § 422.1018 Notice and effect of initial determinations...for the determination, the effect of the determination,...
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.
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.
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.
The Determination of the Orbit Spaces of Compact Coregular Linear Groups
Vittorino Talamini
2015-03-26
Some aspects of phase transitions can be more conveniently studied in the orbit space of the action of the symmetry group. After a brief review of the fundamental ideas of this approach, I shall concentrate on the mathematical aspect and more exactly on the determination of the equations defining the orbit space and its strata. I shall deal only with compact coregular linear groups. The method exposed has been worked out together with prof. G. Sartori and it is based on the solution of a matrix differential equation. Such equation is easily solved if an integrity basis of the group is known. If the integrity basis is unknown one may determine anyway for which degrees of the basic invariants there are solutions to the equation, and in all these cases also find out the explicit form of the solutions. The solutions determine completely the stratification of the orbit spaces. Such calculations have been carried out for 2, 3 and 4-dimensonal orbit spaces. The method is of general validity but the complexity of the calculations rises tremendously with the dimension $q$ of the orbit space. Some induction rules have been found as well. They allow to determine easily most of the solutions for the $(q+1)$-dimensional case once the solutions for the $q$-dimensional case are known. The method exposed is interesting because it allows to determine the orbit spaces without using any specific knowledge of group structure and integrity basis and evidences a certain hidden and yet unknown link with group theory and invariant theory.
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.
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.
Precise orbit determination of a maneuvered GEO satellite using CAPS ranging data
NASA Astrophysics Data System (ADS)
Huang, Yong; Hu, Xiaogong; Huang, Cheng; Yang, Qiangwen; Jiao, Wenhai
2009-03-01
Wheel-off-loadings and orbital maneuvers of the GEO satellite result in additional accelerations to the satellite itself. Complex and difficult to model, these time varying accelerations are an important error source of precise orbit determination (POD). In most POD practices, only non-maneuver orbital arcs are treated. However, for some applications such as satellite navigation RDSS services, uninterrupted orbital ephemeris is demanded, requiring the development of POD strategies to be processed both during and after an orbital maneuver. We in this paper study the POD for a maneuvered GEO satellite, using high precision and high sampling rate ranging data obtained with Chinese Area Positioning System (CAPS). The strategy of long arc POD including maneuver arcs is studied by using telemetry data to model the maneuver thrust process. Combining the thrust and other orbital perturbations, a long arc of 6 days’ CAPS ranging data is analyzed. If the telemetry data are not available or contain significant errors, attempts are made to estimate thrusting parameters using CAPS ranging data in the POD as an alternative to properly account for the maneuver. Two strategies achieve reasonably good data fitting level in the tested arc with the maximal position difference being about 20 m.
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.
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 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. 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 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. PMID:24733025
Orbit Determination Analysis for a Joint UK-Australian Space Surveillance Experiment
NASA Astrophysics Data System (ADS)
Rutten, M.; Harwood, N.; Bennett, J.; Donnelly, P.; Ash, A.; Eastment, J.; Ladd, D.; Gordon, N.; Bessell, T.; Smith, C.; Ritchie, I.
2014-09-01
In February 2014 the UK and Australia carried out a joint space surveillance target tracking, cueing, and sensor data fusion experiment involving the STFC Chilbolton Observatory radar in the UK, the EOS laser-ranging system in Australia and a small telescope operated by DSTO, also in Australia. The experiment, coordinated by DSTL (UK) and DSTO (Aus), was designed to explore the combination of several different, geographically separated sensors for space situational awareness. The primary goal of the experiment was to use data from the radar in the UK to generate an orbital cue to the EOS SLR. A variety of targets sizes and orbits were chosen, under the limitations of observability by both the radar and EOS SLR, in order to explore the variation of cueing accuracy with amount of data incorporated and timeliness from generation. As a secondary objective the effect on cue accuracy of targets in lower orbital regimes was examined. This paper examines the orbit determination techniques used to generate cues from radar and the refined orbits resulting from accumulating SLR data. The construction of tracks using data from all three sensors is explored. Analysis of the accuracy of the orbital reconstructions is made based on comparisons with the measured data and accurate ephemerides provided by the ILRS. The accuracy is tested against the cueing precision requirements for each sensor. Two companion papers describe the experimental goals, execution and achievements (Harwood et. al.) and the sensor aspects of the experiment (Eastment et al.).
The Determination of an Intermediate Perturbed Orbit from Two Position Vectors
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2003-05-01
Based on the theory of intermediate orbits developed earlier by the author of this paper, a new approach is proposed to the solution of the problem of finding the orbit of a celestial body with the use of two position vectors of this body and the corresponding time interval. This approach makes it possible to take into account the main part of perturbations. The orbit is constructed, the motion along which is a combination of two motions: the uniform motion along a straight line of a fictitious attracting center, whose mass varies according to the first Meshchersky law, and the motion around this center. The latter is described by the equations of the Gylden-Meshchersky problem. The parameters of the constructed orbit are chosen so that their limiting values at any reference epoch determine a superosculating intermediate orbit with third-order tangency. The accuracy of approximation of the perturbed motion by the orbits calculated by the classical Gauss method and the new method is illustrated by an example of the motion of the unusual minor planet 1566 Icarus. Comparison of the results obtained shows that the new method has obvious advantages over the Gauss method. These advantages are especially prominent in cases where the angular distances between the reference positions are small.
NASA Astrophysics Data System (ADS)
Blewitt, G.; Kreemer, C.; Hammond, W. C.; Plag, H.-P.; Stein, S.; Okal, E.
The 26 December 2004 Sumatra earthquake Mw 9 2-9 3 generated the most deadly tsunami in history Yet within the first hour the true danger of a major oceanwide tsunami was not indicated by seismic magnitude estimates which were far too low Mw 8 0-8 5 This problem relates to the inherent saturation of early seismic-wave methods Here we show that the earthquake s true size and tsunami potential can be determined using Global Positioning System GPS data up to only 15 minutes after earthquake initiation by tracking the mean displacement of the Earth s surface associated with the arrival of seismic waves Within minutes displacements of 10 mm are detectable as far away as India consistent with results using weeks of data after the event These displacements imply Mw 9 0 - 0 1 indicating a high tsunami potential This suggests existing GPS infrastructure could be developed into an effective component of tsunami warning systems An important aspect for the design of such an envisioned system is real-time access to precise GPS orbit information As a best-case benchmark we solved for satellite orbit and clock parameters simultaneously with Earth rotation and tropospheric refraction using only data up to 20 minutes after the event We then assess the accuracy of rapid earthquake magnitude estimation and hence effectiveness for tsunami warning systems as we vary the quality of GPS orbit information We compare solutions using various strategies including real-time orbit estimation ultra-rapid orbit products from the International GNSS Service
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.
NASA Astrophysics Data System (ADS)
Mann, I. R.
2009-12-01
The Outer Radiation Belt Injection, Transport, Acceleration and Loss Satellite (ORBITALS) mission is proposed as a Canadian Space Agency satellite mission contribution to ILWS. The ORBITALS will provide a unique view of the largely previously unexplored inner magnetosphere. Its mission goal to “understand the acceleration, global distribution, and variability of energetic electrons and ions in the inner magnetosphere” is perfectly aligned with the top geospace priority for the LWS and ILWS programs. This talk will review the ORBITALS scientific objectives and approach to science closure in the context of an international inner magnetosphere constellation, and related ground-based and modelling initiatives. In a 12 hour low inclination orbit, the ORBITALS will come into once daily apogee conjunctions with the extensive ground-based Canadian Geospace Monitoring (CGSM) instrumentation as well as with GOES East and West. Baseline raised perigee will provide both long outer radiation belt dwell times as well as coverage of the outer-most inner radiation belt. In combination, the ORBITALS-CGSM-GOES conjunctions will provide a unique data set with which to address fundamental radiation belt science questions, such as the competition between ULF and VLF acceleration processes, the role of EMIC and VLF waves in loss, and the relationship between these processes and plasmaspheric cold plasma dynamics. As an example, radial diffusion rates derived from GOES and THEMIS electric and magnetic field power will be compared to previous empirical approaches, illustrating the importance of global measurements for identifying dominant or active acceleration mechanisms. In combination with the approved NASA LWS RBSP mission, and the proposed Japanese ERG satellite, the ORBITALS-RBSP-ERG three petal constellation will resolve the spatio-temporal ambiguities and global dynamics and morphology of the Earths radiation belts.
From Astrometry to Celestial Mechanics: Orbit Determination with Very Short Arcs
NASA Astrophysics Data System (ADS)
Milani, Andrea; Kneževi?, Zoran
2005-04-01
Contemporary surveys provide a huge number of detections of small solar system bodies, mostly asteroids. Typically, the reported astrometry is not enough to compute an orbit and/or perform an identification with an already discovered object. The classical methods for preliminary orbit determination fail in such cases: a new approach is necessary. When the observations are not enough to compute an orbit we represent the data with an attributable (two angles and their time derivatives). The undetermined variables range and range rate span an admissible region of solar system orbits, which can be sampled by a set of Virtual Asteroids (VAs) selected by an optimal triangulation. The attributable results from a fit and has an uncertainty represented by a covariance matrix, thus the predictions of future observations can be described by a quasi-product structure (admissible region times confidence ellipsoid), which can be approximated by a triangulation with each node surrounded by a confidence ellipsoid. The problem of identifying two independent short arcs of observations has been solved. For each VA in the admissible region of the first arc we consider prediction at the time of the second arc and the corresponding covariance matrix, and we compare them with the attributable of the second arc with its own covariance. By using the penalty (increase in the sum of squares, as in the algorithms for identification) we select the VAs which can fit together both arcs and compute a preliminary orbit. Even two attributables may not be enough to compute an orbit with a convergent differential corrections algorithm. The preliminary orbits are used as first guess for constrained differential corrections, providing solutions along the Line Of Variations (LOV) which can be used as second generation VAs to further predict the observations at the time of a third arc. In general the identification with a third arc will ensure a least squares orbit, with uncertainty described by the covariance matrix.
Mitigation of ionospheric scintillation effects in kinematic LEO precise orbit determination
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-Gürr, 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 with the DORIS/SPOT2 system: First results
NASA Astrophysics Data System (ADS)
Nouel, F.; Berthias, J. P.; Broca, P.; Comps, A.; Deleuze, M.; Guitart, A.; Jayles, C.; Laudet, P.; Pierret, C.; Piuzzi, A.
1991-12-01
One of the main challenges of the French/U.S. oceanographic mission, TOPEX/Poseidon, is the need to determine the position of the spacecraft with an accuracy of 10 cm in the radial direction after 10 days. In order to reach this level of accuracy, a tracking system, Doppler Orbitography and Radiopositioning Integrated by Satellite (DORIS), was developed. This system is composed of a network of about 40 dual frequency beacons scattered all over the world and a receiver on board a spacecraft. The spaceborne processor integrates the Doppler shifts over 10 second intervals and downloads the results and the measurements time tags to the ground control center. An orbit determination software, ZOOM, was developed to generate precise orbits. This program has the capability to process the DORIS data type in addition to all other conventional measurements. As a proof of concept and feasibility test, the first DORIS receiver was placed on the SPOT 2 Earth observation satellite. The whole system (hardware and software) has worked successfully since Feb. 1990, enabling the routine computation of the orbit of SPOT 2 for more than one year. An overview of the orbit production process including the operational set up and the dynamical and measurement models is presented. The orbit quality assessment procedures are described and the ability to ensure the consistency of the system in spite of the large number of independent oscillators, is demonstrated.
NASA Astrophysics Data System (ADS)
Sakamoto, Yuji; Yoneyama, Akari
Recently the activities for the development of micro- and nano- satellites like a cubesat are being increased. A lot of satellites are operated at a ground station located in a university, but public orbit information is used. In this paper, the orbit determination system is constructed using observations of Doppler frequency at a low-cost ground station. The traditional batch state estimation filter is used, and the orbital elements of satellite position and velocity in inertia coordinates are determined from the Doppler frequency of receiving signals. The verification tests are conducted using UHF-band signals (about 430 MHz) received at an amateur radio station and S-band signals (about 2.2 GHz) received at a 2.4-m parabola antenna from real satellites, and the valid performance is conformed for tracking satellites not depending on public orbit information. The error is max. 0.3 degrees in direction and max. 3.9 km in position for UHF-band signals, and max. 1.1 degrees in direction and max. 37 km in position for S-band signals. This method is valid especially for low-earth-orbit satellites with large Doppler effect.
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).
Arrival and departure impulsive Delta V determination for precessing Mars parking orbits
NASA Technical Reports Server (NTRS)
Desai, Prasun N.; Buglia, James J.
1992-01-01
An attempt is made to develop a method for realistic estimation of the initial LEO mass. The method takes into account the actual geometry between the inbound and outbound hyperbolic asymptotes and the parking orbit, along with precession effects caused by the oblateness of Mars, in calculating the arrival and departure Delta V values. Three mission scenarios alternative to the arbitrarily assumed two tangential periapsis burns are described: a tangential periapsis arrival and an in-plane departure; an in-plane arrival and in-plane departure; and a tangential periapsis arrival and a 3D departure. Results obtained by the method under consideration compared well with a trajectory integration code, where the differences in the initial LEO orbit mass were within one percent, for all three cases. The method is considered to be an ideal tool for preliminary mission design, since it reduces the analysis computation time with minimal loss in accuracy.
Hubble and Planck scale limits on the determination of orbital angular momentum states of light
F. Tamburini; B. Thidé; A. Sponselli
2012-01-16
We review Heisenberg's uncertainty principle for the orbital angular momentum (OAM) of light. By taking into account the largest and smallest scales present in nature, such as the the Hubble radius and the Planck length, we have found that there exist upper and lower physical limits to the determination of the OAM of a photon.
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.
Real-time, autonomous precise satellite orbit determination using the global positioning system
David Ben Goldstein
2000-01-01
The desire for autonomously generated, rapidly available, and highly accurate satellite ephemeris is growing with the proliferation of constellations of satellites and the cost and overhead of ground tracking resources. Autonomous Orbit Determination (OD) may be done on the ground in a post-processing mode or in real-time on board a satellite and may be accomplished days, hours or immediately after
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2011-07-01
Transformations of differential equations of the methods for determining the Lyapunov Characteristic Indicator and MEGNO indicators are suggested. The transformations improve the behavior of the differential equations by their simultaneous numerical integration. The use of the transformed equations is especially efficient for the investigation of orbits in stochastic regimes.
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.
Lunar Reconnaissance Orbiter Camera Narrow Angle Cameras: Laboratory and Initial Flight Calibration
D. C. Humm; M. Tschimmel; B. W. Denevi; S. Lawrence; P. Mahanti; T. N. Tran; P. C. Thomas; E. Eliason; M. S. Robinson
2009-01-01
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
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 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.
Determination of kinematic state of an orbiting multibody using GNSS signals
NASA Astrophysics Data System (ADS)
Palmerini, Giovanni B.; Gasbarri, Paolo; Toglia, Chiara
2009-06-01
Precise attitude determination of the members of a free-flying multibody system is a not so immediate task, due essentially to the large motion of its appendages coupled with their relevant flexibility effects. In fact, sensors used to this aim in current projects, such as optical encoders usually positioned near the joints of each arm, are almost blind to these effects, and clusters of specific redundant sensors should, therefore, be required in order to reconstruct both elastic deformations and rigid motion. Satellite navigation systems (GNSS) offer a suitable and reliable solution to this problem. To exploit the phase of the signal, instead of the traditional pseudo random code, ensures a very high accuracy of the order of magnitude of centimeter. Such a process requires the solution of an initial ambiguity problem, related to the number of integer wavelength included in the length of the member. The aim of the paper is to investigate the capability of this GNSS based technique to reconstruct the kinematics of a flexible multibody system orbiting around the Earth. This analysis requires a simulation including both the multibody dynamics and the navigation system constellation to define the satellites lines-of-sight at each time step. Concerning multibody equations of motion, a Newtonian formulation is adopted in this work. A special attention is required about the choice of the state variables. As the internal forces are associated to the relative displacements between the bodies, which are small fractions of the distance of the multibody spacecraft from the center of the Earth, the task of obtaining these forces from inertial coordinates could be impossible from a numerical point of view. So, the problem is reformulated in such a way that the equation of motion of the system contains global equations, with no internal forces, and local equations, with internal forces. In the latter, only quantities of the same order of the spacecraft dimensions are present. Accuracies achievable in LEO orbit with current GPS and upcoming Galileo systems are evaluated to show the interest of the proposed technique.
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.
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.
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.
Determination of the area and mass distribution of orbital debris fragments
NASA Technical Reports Server (NTRS)
Badhwar, Gautam D.; Anz-Meador, Phillip D.
1989-01-01
A technique is described to estimate the area-to-mass ratio of debris fragments using orbital fragments obtained by radar. The area-to-mass ratio of about 2600 fragments arising from the breakup of 24 artificial satellites was determined; an analysis of the data on about 200 objects with known mass, size, and shape has been made, and a calibration of the observed radar cross-section (RCS) to the effective area of these objects has provided a method to estimate the effective area of debris fragments. From the knowledge of the effective area and the estimated area-to-mass ratio, the mass and area distribution of each of the known breakup has been obtained. As a function of time, the orbital elements can be used to invert any propagation algorithm to yield the area-to-mass ratio of an orbiting object.
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 Astrophysics Data System (ADS)
Porfilio, M.; Piergentili, F.; Graziani, F.
In September 2003 the Group of Astrodynamics of the University of Rome ``La Sapienza'' (GAUSS) carried out a two-site observation campaign devoted to the autonomous orbit determination of objects in the geosynchronous region. Two 40 cm aperture Ritchey-Chrétien devices were employed: the f/7.5 ``Collepardo Automatic Telescope'' (CAT, located in Collepardo, Italy) and a f/5 tube of the ``Observatori Astronòmic de Mallorca'' (OAM, located in Mallorca, Spain). The baseline between the sites is 916 km. 3 s long, 1 minute apart exposures were simultaneously taken in sidereal tracking mode, looking at the same arcs of the GEO ring; the fields of view allowed to see a few satellites in two successive frames from both sites, thus providing two positions: the Lambert theorem has been exploited to determining the orbits. A first order approximation of the targets angular motion has been used to fix synchronism errors. Of course, the longer the time interval between positions, the lower the effect of measurements errors. Nevertheless, the only way to have quite distant points would be tracking the satellite, which is typically not suitable for a surveillance campaign, thus not interesting from a practical standpoint. Currently, in the Measurement Working Group of the Inter-Agency Space Debris Co-ordination Committee (IADC), the orbits of the objects detected during GEO optical observation campaigns, are estimated under the assumption of null eccentricity. This is the only way, if one telescope is used and if only a few observations are available. Obviously, the hypothesis of circular orbit provides excellent results for actually geostationary satellites and definitely incorrect estimates for high eccentricity objects. The systematic cooperation of couples of observatories, would provide good orbit determination, for instance, for GTO debris. In the paper the results of the orbit determination from the September 2003 campaign are reported. More in detail, the outcomes of some classical methods for solving the Lambert theorem, are compared with the least squares improved solutions, with the circular orbit assumption results and with the TLEs.
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.
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.
Topex/Jason combined GPS/DORIS orbit determination in the tandem phase
NASA Astrophysics Data System (ADS)
Willis, P.; Haines, B.; Bar-Sever, Y.; Bertiger, W.; Muellerschoen, R.; Kuang, D.
In December 2001, the Jason-1 satellite was launched to extend the long-term success of the TOPEX/POSEIDON (T/P) oceanographic mission. The goals for the Jason-1 mission represent both a significant challenge and rare opportunity for precise orbit determination (POD) groups. Like its predecessor, Jason-1 carries three types of POD systems : a GPS receiver, a DORIS receiver and a laser retro-reflector. In view of the 1-cm goal for radial orbit accuracy, several major improvements have been made to the POD systems: 1) the GPS TurboRogue Space Receiver (TRSR) tracks up to 12 GPS spacecraft using advanced codeless tracking techniques; 2) a newly developed DORIS receiver can track two ground beacons simultaneously with lower noise. In addition, the satellite itself features more straightforward attitude behavior, and a symmetric shape, simplifying the orbit determination models compared to T/P. On the other hand, the area-to-mass ratio for Jason-1 is larger, implying larger potential surface-force errors. This paper will present Jason-1 POD results obtained at JPL using the Gipsy-Oasis II (GOA). Results from standard tests (orbit overlaps, Laser control points, altimeter crossovers) suggest that 1 to 2 cm radial orbit precision is already being achieved using a reduced-dynamic filter approach. New DORIS POD strategies will be an emphasis of this paper. These strategies make full profit of the additional number of common DORIS observations due to the T/P-Jason-1 tandem mode of orbit as well the additional dual-channel capability of the upgraded JASON receiver (allowing simultaneous tracking of two ground stations). New information on the satellite's time scale is availed through this new filtering strategy. Results show that a slight improvement could be gained on DORIS-based orbits using this strategy. This improvement may become more evident in the near future, as new launches will bring to 6 the total number of satellites collecting DORIS observations on the same day. Building on these results, we have extended the GOA software capability to more fully exploit the combined benefit of both GPS and DORIS measurements from T/P and Jason-1 in their preliminary tandem mode. POD test results will be used to demonstrate the accuracy of these orbits and to compare results in different cases: GPS-alone, DORIS-alone, and GPS and DORIS together in both single- and multi-satellite modes. Finally, plans for future software enhancements, processing strategies and modeling improvements will be presented.
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 km× 100 km lunar orbit, the position errors are better than 30 meters, and for the CE-2's 15 km× 100 km orbit, the position errors are better than 45 meters. The observational data with the delta differential one-way ranging (? DOR) from the CE-2's X-band monitoring and control system experimental are analyzed. It is concluded that the accuracy of ? 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 km× 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 (100× 100), but for the 15 km× 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
Precise GLONASS orbit determination within the IGS/IGLOS Pilot Project
NASA Astrophysics Data System (ADS)
Weber, R.; Fragner, E.; Slater, J. A.; Habrich, H.; Glotov, V.; Romero, I.; Schaer, S.
During the past 3.5 years, Russia has launched a number of new GLONASS satellites to renew step by step the whole space segment. 10 satellites are currently operational, while the new GLONASS-M satellite is still in commission phase. In parallel, the International GPS Service (IGS) established in 2000 a GLONASS Service Pilot Project (IGLOS) as a follow-on to the very successful International GLONASS Experiment (IGEX-98) that ended in April 1999. The IGLOS Pilot Project has a global network of over 40 tracking stations with dual-frequency GLONASS receivers, collocated with IGS GPS stations. As with the GPS data, the GLONASS data are collected continuously and archived in RINEX format at the IGS Global Data Centers. In addition, three of the satellites are routinely tracked by the International Laser Ranging Service, taking advantage of the retro-reflector array carried by each satellite. Three organizations, ESA/ESOC (European Space Operations Center, Germany), BKG (Bundesamt f. Kartographie u. Geodäsie, Germany) and CODE (University Berne, Switzerland), compute precise orbits from the receiver tracking data, while the Russian Mission Control Center computes orbits from the laser ranging data. The laser data provide a means to validate orbit determination modeling and to investigate orbit residuals. These 4 independent orbits, which are consistent at the 10-15cm level, are used by the IGLOS Analysis Coordinator to compute one combined IGLOS orbit for each satellite. The IGLOS project is a working model of a 2-system GNSS. Time standardization, reference frames and file formats have been addressed to handle GPS and GLONASS data in the same operations. The combined systems contain over 36 satellites, which can be exploited for many applications. In this context time transfer, real-time navigation, and atmospheric studies are valid examples which can benefit from the additional data that GLONASS provides. Finally, taking into consideration the upcoming GALILEO system, IGLOS can pave the way for handling data from a multi-system GNSS.
NASA Astrophysics Data System (ADS)
Shang, Lin; Liu, Guohua; Zhang, Rui; Li, Guotong
2013-04-01
This paper focuses on the information fusion problem of integrated autonomous orbit determination using the observations from inter-satellite-link (ISL), X-ray pulsars and star sensors. A step Kalman filter structure is proposed to solve the information fusion problem of multiple subsystems that have greatly different filtering precision. The subsystems are grouped according to their measurement accuracy and the state parameters and covariance matrix of a group can be calculated using the federated filter structure and propagated to the next group step-by-step. Simulation results show that the mean user range error (URE) of the constellation will be less than 1.5 m in 60 days using the step Kalman filter structure for information fusion. And it has better performance than the federated structure in dealing with information fusion of the astronomical observations and the ISL ranging measurements in integrated autonomous orbit determination.
Application of the total least squares method to the determination of preliminary orbit
NASA Astrophysics Data System (ADS)
Chen, Wu-shen; Zhang, Jing; Ma, Jing-yuan; Lu, Ben-kui
2006-10-01
Based on the analysis of the characteristics of the equations of condition in the UVM2 (Unit Vector Method 2), the total least squares method (TLS) is introduced into the orbital determination and the linearization of the vis-viva formula in the original algorithm is thereby avoided. The calculated results from simulation and observation data show that the application of TLS to UVM2 is effective.
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.
20 CFR 663.515 - What is the process for initial determination of provider eligibility?
Code of Federal Regulations, 2010 CFR
2010-04-01
...process for initial determination of provider eligibility? 663.515 Section...INVESTMENT ACT Eligible Training Providers § 663.515 What is the process for initial determination of provider eligibility? (a) To be...
49 CFR 381.405 - Who determines whether a pilot program should be initiated?
Code of Federal Regulations, 2010 CFR
2010-10-01
...2010-10-01 false Who determines whether a pilot program should be initiated? 381.405 Section...SAFETY REGULATIONS WAIVERS, EXEMPTIONS, AND PILOT PROGRAMS Initiation of Pilot Programs § 381.405 Who determines...
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.
A demonstration of unified TDRS/GPS tracking and orbit determination
NASA Astrophysics Data System (ADS)
Haines, B.; Lichten, S.; Srinivasan, J.; Young, L.
1995-05-01
We describe results from an experiment in which TDRS and GPS satellites were tracked simultaneously from a small (3 station) ground network in the western United States. We refer to this technique as 'GPS-like tracking' (GLT) since the user satellite - in this case TDRS - is essentially treated as a participant in the GPS constellation. In the experiment, the TDRS K(sub space-to-ground link (SGL) was tracked together with GPS L-band signals in enhanced geodetic-quality GPS receivers (TurboRogue). The enhanced receivers simultaneously measured and recorded both the TDRS SGL and the GPS carrier phases with sub-mm precision, enabling subsequent precise TDRS orbit determination with differential GPS techniques. A small number of calibrated ranging points from routine operations at the TDRS ground station (White Sands, NM) were used to supplement the GLT measurements in order to improve determination of the TDRS longitude. Various tests performed on TDRS ephemerides derived from data collected during this demonstration - including comparisons with the operational precise orbit generated by NASA Goddard Space Flight Center - provide evidence that the TDRS orbits have been determined to better than 25 m with the GLT technique.
A new approach for precise orbit determination based on raw GNSS measurements
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-Gürr, Torsten
2013-04-01
Kinematic orbit determination based on GNSS measurements is a core element for gravity field determination from Low Earth Orbiting Satellites as for example GRACE and GOCE. Presently used algorithms for kinematic orbit determination are based on observation combinations, like for example the ionosphere-free combination or double differences. In this presentation we will introduce a new approach which is based on raw measurements. The method uses all available observations directly without forming any kind of differences or linear combinations. This includes phase as well as code observations. Every observation is represented by a separate observation equation in a least squares adjustment. All errors and influences which are known with sufficient accuracy are corrected beforehand, otherwise they are added to the estimation as additional parameters. This enables a very flexible handling of the different errors included in the observations. For example antenna center variations can be simply added to the estimation as additional parameters. This method is also well suited for future developments, as new observation types can be included by just adding a new observation equation to the design matrix. Due to the fact that the observations are used directly, the integer ambiguities are accessible even when processing a single receiver. Results based on GRACE and GOCE data will be presented.
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
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.
Determination Of The Orbit Of The Planetary Companion To The Metal Rich Star HD 45350
Michael Endl; William D. Cochran; Robert A. Wittenmyer; Artie P. Hatzes
2006-02-28
We present the precise radial velocity (RV) 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 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 semi-amplitude K of 57.4 m/s. The planet has a minimum mass of m sin i = 1.82 +- 0.14 M_Jup and an orbital semi-major axis of a = 1.92 +-0.07 AU.
NASA Technical Reports Server (NTRS)
Trujillo, B. M.
1986-01-01
This paper presents the technique and results of maximum likelihood estimation used to determine lift and drag characteristics of the Space Shuttle Orbiter. Maximum likelihood estimation uses measurable parameters to estimate nonmeasurable parameters. The nonmeasurable parameters for this case are elements of a nonlinear, dynamic model of the orbiter. The estimated parameters are used to evaluate a cost function that computes the differences between the measured and estimated longitudinal parameters. The case presented is a dynamic analysis. This places less restriction on pitching motion and can provide additional information about the orbiter such as lift and drag characteristics at conditions other than trim, instrument biases, and pitching moment characteristics. In addition, an output of the analysis is an estimate of the values for the individual components of lift and drag that contribute to the total lift and drag. The results show that maximum likelihood estimation is a useful tool for analysis of Space Shuttle Orbiter performance and is also applicable to parameter analysis of other types of aircraft.
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.
NASA Technical Reports Server (NTRS)
Shanklin, R. E., Jr.; Lee, T.; Samii, M.; Mallick, M. K.; Cappellari, J. O., Jr.
1982-01-01
The results of a comparative orbit determination study of four global atmospheric density models (modified Harris-Priester, Jacchia-Roberts, Mass Spectrometer/Incoherent Scatter (MSIS), and Simple Exponential Model (SEM)) are presented. Utilizing these models, definitive orbit determination consistency and accuracy are evaluated using the maximum position differences that occur during 6-hour overlap periods between ephemerides generated from 30-hour data arcs. Propagated ephemerides are compared with definitive orbit solutions to evaluate predictive accuracy. The results indicate that, for satellites above 300 kilometers, all four atmospheric density models produce comparable orbit determination accuracies when an atmospheric drag scaling factor and the satellite state vector are estimated in the orbit determination process.
Federal Register 2010, 2011, 2012, 2013, 2014
2013-07-25
...INTERNATIONAL TRADE COMMISSION [Investigation No. 337-TA-869] Certain Robotic Toys and Components Thereof; Commission Determination Not To Review an Initial Determination Granting a Joint Motion for...
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.
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.
20 CFR 418.1310 - When may you request that we make a new initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
...2010-04-01 false When may you request that we make a new initial determination? 418... § 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...
Preliminary Orbit Determination Using the Gauss and the Double-R Iteration Methods
NASA Astrophysics Data System (ADS)
Mahooti, M.
2009-04-01
In order to obtain orbital elements of a satellite from angular measurements, there are three approaches: Laplace's, Gauss's, and double r-iteration. The Gauss and the double r-iteration techniques are commonly used for practical purposes. These methods use three sets of chronologically ordered gimbal angle measurements from up to three separate tracking stations to determine the Cartesian components of position and velocity. The angle data set can be distributed over an orbital arc of less than 60 degrees in mean anomaly for the Gauss method and up to 360 degrees in mean anomaly for the double d-iteration method. The epoch for the position and velocity corresponds to the time of the second measurement set. The methods are deterministic since the six measurement components yield the six position and velocity components. Description and test of these two approaches with the STK are the aims of this paper. Keywords: Preliminary orbit determination, Gauss method, Double r-iteration method, Angular observations
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)
Paetzold, M.; Eidel, W.; Haeusler, B.; Schmitt, J.
One of the science objectives of the Rosetta Radio Science Investigations (RSI) experiment is the determination of the total cometary mass flux (gas and dust) onto the Rosetta spacecraft when in orbit about the nucleus of comet Wirtanen, starting in 2012. The RSI experiment will use the spacecrafts radio carrier frequencies at X-band (8.4 GHz) and S-band (2.3 GHz) in order to measure slight changes in the relative velocity between the spacecraft and the ground station on Earth (Doppler effect) induced by perturbing forces, mainly the gas and dust flow onto the spacecraft. The cometary mass flux is estimated based on the gas and dust production rates (3 AU to perihelion) observed at the last Wirtanen apparition. The gas flow will be the dominant perturber of the Rosetta orbit and the force acting on the spacecraft will exceed the gravity attraction of the nucleus if the nucleus is within two astronomical units heliocentric distance.
Determination of The Cometary Mass Flux Onto The Rosetta Spacecraft When In Orbit About Wirtanen
NASA Astrophysics Data System (ADS)
Pätzold, M.; Häusler, B.; Schmitt, J.; Wennmacher, A.
One of the science objectives of the Rosetta Radio Science Investigation (RSI) ex- periment is the determination of the total cometary mass flux (gas and dust) onto the Rosetta spacecraft when in orbit about the nucleus of comet Wirtanen starting in 2012. The RSI experiment will use the spacecrafts radio carrier frequencies at X-band (8.4 GHz) and S-band (2.3 GHz) in order to measure slight changes in the relative velocity between the spacecraft and the ground station on Earth (Doppler effect) induced by the perturbing force of the cometary gas and dust flow onto the spacecraft. These per- turbing force is estimated based on the observed gas and dust production rates (3 AU to perihelion) from the last Wirtanen apparition. The gas flow will be the dominant perturber of the spacecraft orbit, the force will exceed even the gravity attraction of the nucleus if the comet is within two astronomical units heliocentric distance.
Magnetospheric plasma analyzer: Initial three-spacecraft observations from geosynchronous orbit
McComas, D.J.; Bame, S.J.; Barraclough, B.L.; Donart, J.R.; Elphic, R.C.; Gosling, J.T.; Moldwin, M.B.; Moore, K.R.; Thomsen, M.F.
1993-08-01
The first three magnetospheric plasma analyzer (MPA) instruments have been returning data from geosynchronous orbit nearly continuously since late 1989, 1990, and 1991. These identical instruments provide for the first time simultaneous plasma observations from three widely spaced geosynchronous locations. The MPA instruments measure the three-dimensional velocity space distributions of both electrons and ions with energies between 1 eV/q and 40 keV/q. The authors use the simultaneous observations from three longitudinally separated spacecraft to synthesize a synoptic view of the morphology of the magnetosphere at geosynchronous orbit over a 6-week interval in early 1992. The MPA observations indicate that the spacecraft encountered seven regions with characteristic plasma populations during this period: (1) the cool, dense plasmasphere (13.1% of the data); (2) a warmer, less dense plasma trough (22.5%); (3) the hot plasma sheet (40.3%); (4) a combination of plasma trough and plasma sheet (18.6%); (5) an empty trough region, devoid of plasma sheet, plasmasphere, or plasma trough populations (4.3%); (6) the magnetosheath and/or low-latitude boundary layer (0.7%); and (7) the lobe (0.3%). As expected, the plasmapause is found to have a highly variable shape, at various times showing (1) a stable dusk side bulge, (2) a variable bulge which expands, contracts, and moves, (3) an overall expansion and contraction of the plasmasphere, and (4) even more complicated behavior which is best accounted for by large-scale structure of the plasmapause and/or disconnected plasma blobs. During the 6 weeks of data examined, the magnetosheath was encountered on several occasions at synchronous orbit, preferentially on the prenoon side of the magnetosphere. For the first time, simultaneous prenoon and postnoon observations confirm this asymmetry and demonstrate that the magnetopause shape can be highly asymmetric about the Earth-Sun line. 26 refs., 11 figs., 3 tabs.
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.
Y. Chen; R. H. W. Friedel; G. D. Reeves; T. G. Onsager; M. F. Thomsen
2005-01-01
We develop and test a methodology to determine the relativistic electron phase space density distribution in the vicinity of geostationary orbit by making use of the pitch-angle resolved energetic electron data from three Los Alamos National Laboratory geosynchronous Synchronous Orbit Particle Analyzer instruments and magnetic field measurements from two GOES satellites. Owing to the Earth's dipole tilt and drift shell
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2013-12-01
A new method is proposed for computing the preliminary orbit of a small celestial body from three pairs of range and range rate observations. The method is based on using the superosculating intermediate orbit with a fourth-order tangency that we previously constructed. This intermediate orbit allows for most of the perturbations in the motion of the body under study. The methodical error of orbit determination by the proposed method is three orders smaller than the corresponding error of the commonly used approach based on the construction of the unperturbed Keplerian orbit. Using the examples of finding the orbits of artificial Earth satellites, the results obtained by the procedure implementing the traditional approach and the new method are compared. The comparison shows that the new method is a highly efficient means for studying perturbed motion.
NASA Astrophysics Data System (ADS)
Paredes-Gil, Katherine; Jaque, Pablo
2015-01-01
The Rusbnd PR3 bonds of 1-2(a-b)-PC, Rudbnd CHPh bonds of 1a-b, 2-Inact/Act and 1a-b, 2-RCB were analyzed by charge decomposition (CDA) and natural bond orbital (NBO). We have found that the dissociation step of the Rusbnd PR3 bond is driven by charge transfer, while the RCB by polarization effects. Furthermore, the ?(Cipso)-?*(Rudbnd C) interaction was associated with delocalization effects in the benzylidene ring. Likewise, the nature of the rotameric changes in the carbene was studied through the resonance stabilization energy (ENLW). 2 presented a lower ?ENLW (Inactive ? Active) than 1a-b, which confirms that the delocalization effects are related to a low carbene rotameric energy.
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.
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 Astrophysics Data System (ADS)
Pasetto, S.; Grebel, E. K.; Berczik, P.; Chiosi, C.; Spurzem, R.
2011-01-01
We present a new study of the evolution of the Carina dwarf galaxy that includes a simultaneous derivation of its orbit and star formation history. The structure of the galaxy is constrained through orbital parameters derived from the observed distance, proper motions, radial velocity, and star formation history. The different orbits admitted by the large proper motion errors are investigated in relation to the tidal force exerted by an external potential representing the Milky Way. Our analysis is performed with the aid of fully consistent N-body simulations that are able to follow the dynamics and the stellar evolution of the dwarf system in order to determine the star formation history of Carina self-consistently. We also find a star formation history characterized by several bursts, partially matching the observational expectation. We find also compatible results between dynamical projected quantities and the observational constraints. The possibility of a past interaction between Carina and the Magellanic Clouds is also separately considered and deemed unlikely. Appendices are only available in electronic form at http://www.aanda.org
Determination of Orbital Parameters for Visual Binary Stars Using a Fourier-Series Approach
NASA Astrophysics Data System (ADS)
Brown, D. E.; Prager, J. R.; DeLeo, G. G.; McCluskey, G. E., Jr.
2001-12-01
We expand on the Fourier transform method of Monet (ApJ 234, 275, 1979) to infer the orbital parameters of visual binary stars, and we present results for several systems, both simulated and real. Although originally developed to address binary systems observed through at least one complete period, we have extended the method to deal explicitly with cases where the orbital data is less complete. This is especially useful in cases where the period is so long that only a fragment of the orbit has been recorded. We utilize Fourier-series fitting methods appropriate to data sets covering less than one period and containing random measurement errors. In so doing, we address issues of over-determination in fitting the data and the reduction of other deleterious Fourier-series artifacts. We developed our algorithm using the MAPLE mathematical software code, and tested it on numerous "synthetic" systems, and several real binaries, including Xi Boo, 24 Aqr, and Bu 738. This work was supported at Lehigh University by the Delaware Valley Space Grant Consortium and by NSF-REU grant PHY-9820301.
20 CFR 418.3615 - Will we mail you a notice of the initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
...2010-04-01 2010-04-01 false Will we mail you a notice of the initial determination...Review Process § 418.3615 Will we mail you a notice of the initial determination? (a) We will mail a written notice of the initial...
Code of Federal Regulations, 2010 CFR
2010-10-01
...determining eligibility after the initial one or two months. 233.26 Section 233.26...determining eligibility after the initial one or two months. (a) Under retrospective...eligibility, following the initial one or two months under § 233.24, shall be...
Code of Federal Regulations, 2011 CFR
2011-10-01
...determining eligibility after the initial one or two months. 233.26 Section 233.26...determining eligibility after the initial one or two months. (a) Under retrospective...eligibility, following the initial one or two months under § 233.24, shall be...
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.
14 CFR 1206.603 - Procedures and time limits for initial determinations.
Code of Federal Regulations, 2010 CFR
2010-01-01
14 Aeronautics and Space 5 2010-01-01...false Procedures and time limits for initial determinations...603 Aeronautics and Space NATIONAL AERONAUTICS AND SPACE ADMINISTRATION AVAILABILITY...603 Procedures and time limits for initial...
14 CFR 1206.603 - Procedures and time limits for initial determinations.
Code of Federal Regulations, 2011 CFR
2011-01-01
14 Aeronautics and Space 5 2011-01-01...true Procedures and time limits for initial determinations...603 Aeronautics and Space NATIONAL AERONAUTICS AND SPACE ADMINISTRATION AVAILABILITY...603 Procedures and time limits for initial...
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.
Kliore, A J; Patel, I R; Nagy, A F; Cravens, T E; Gombosi, T I
1979-07-01
Pioneer Venus orbiter dual-frequency radio occultation measurements have produced many electron density profiles of the nightside ionosphere of Venus. Thirty-six of these profiles, measured at solar zenith angles (chi) from 90.60 degrees to 163.5 degrees , are discussed here. In the "deep" nightside ionosphere (chi > 110 degrees ), the structure and magnitude of the ionization peak are highly variable; the mean peak electron density is 16,700 +/- 7,200 (standard deviation) per cubic centimeter. In contrast, the altitude of the peak remains fairly constant with a mean of 142.2 +/- 4.1 kilometers, virtually identical to the altitude of the main peak of the dayside terminator ionosphere. The variations in the peak ionization are not directly related to contemporal variations in the solar wind speed. It is shown that electron density distributions similar to those observed in both magnitude and structure can be produced by the precipitation on the nightside of Venus of electron fluxes of about 108 per square centimeter per second with energies less than 100 electron volts. This mechanism could very likely be responsible for the maintenance of the persistent nightside ionosphere of Venus, although transport processes may also be important. PMID:17778916
NASA Technical Reports Server (NTRS)
Chato, David J.
1991-01-01
The results are presented of a series of no-vent fill experiments conducted on a 175 cu ft flightweight hydrogen tank. The experiments consisted of the nonvented fill of the tankage with liquid hydrogen using two different inlet systems (top spray, and bottom spray) at different tank initial conditions and inflow rates. Nine tests were completed of which six filled in excess of 94 percent. The experiments demonstrated a consistent and repeatable ability to fill the tank in excess of 94 percent using the nonvented fill technique. Ninety-four percent was established as the high level cutoff due to requirements for some tank ullage to prevent rapid tank pressure rise which occurs in a tank filled entirely with liquid. The best fill was terminated at 94 percent full with a tank internal pressure less than 26 psia. Although the baseline initial tank wall temperature criteria was that all portions of the tank wall be less than 40 R, fills were achieved with initial wall temperatures as high as 227 R.
PSA: A program to streamline orbit determination for launch support operations
NASA Technical Reports Server (NTRS)
Legerton, V. N.; Mottinger, N. A.
1988-01-01
An interactive, menu driven computer program was written to streamline the orbit determination process during the critical launch support phase of a mission. Residing on a virtual memory minicomputer, this program retains the quantities in-core needed to obtain a least squares estimate of the spacecraft trajectory with interactive displays to assist in rapid radio metric data evaluation. Menu-driven displays allow real time filter and data strategy development. Graphical and tabular displays can be sent to a laser printer for analysis without exiting the program. Products generated by this program feed back to the main orbit determination program in order to further refine the estimate of the trajectory. The final estimate provides a spacecraft ephemeris which is transmitted to the mission control center and used for antenna pointing and frequency predict generation by the Deep Space Network. The development and implementation process of this program differs from that used for most other navigation software by allowing the users to check important operating features during development and have changes made as needed.
Galileo precise orbit and clocks determination at the CNES-CLS IGS Analysis Center
NASA Astrophysics Data System (ADS)
Loyer, S.; Mercier, F.; Capdeville, H.; Andrianavonimiarina, J.; Mezerette, A.; Perosanz, F.; Boulanger, C.; Lestarquit, L.
2013-12-01
Thanks to the IGS (International GNSS Service) Multi-GNSS Experiment (M-GEX), signals from new GNSS satellites like Galileo are now available. CNES and IGN joined their efforts to contribute with others international agency to the densification of this multi-GNSS global network through the REGINA project. Since mid-2012 we process the data from the global M-GEX tracking network (~50 stations) to determine precise orbit of the Galileo satellites. The strategy followed at the CNES-CLS IGS Analysis Center uses a combined computation of GPS (E1/E2) and GALILEO (E1/E5a) data. This strategy allows the continuous determination of receiver clocks and the recovering of pseudo-range and phase biases between receivers and between GPS and Galileo observations. In parallel with orbit and clocks solutions we provide a daily monitoring of the network receivers coordinates. The quality of our products is accessed through their comparison with the other agencies participating to the M-GEX project as well as PPP (Precise Point Positioning) tests.
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.
Code of Federal Regulations, 2010 CFR
2010-07-01
...the following ways: (1) If 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 those unfunded...
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.
Code of Federal Regulations, 2013 CFR
2013-04-01
...initial or reconsidered determination made by us ever appropriate? 418.1345 Section...initial or reconsidered determination made by us ever appropriate? We may reopen an initial or reconsidered determination made by us when the conditions for reopening are...
Code of Federal Regulations, 2011 CFR
2011-04-01
...initial or reconsidered determination made by us ever appropriate? 418.1345 Section...initial or reconsidered determination made by us ever appropriate? We may reopen an initial or reconsidered determination made by us when the conditions for reopening are...
Code of Federal Regulations, 2014 CFR
2014-04-01
...initial or reconsidered determination made by us ever appropriate? 418.1345 Section...initial or reconsidered determination made by us ever appropriate? We may reopen an initial or reconsidered determination made by us when the conditions for reopening are...
Code of Federal Regulations, 2010 CFR
2010-04-01
...initial or reconsidered determination made by us ever appropriate? 418.1345 Section...initial or reconsidered determination made by us ever appropriate? We may reopen an initial or reconsidered determination made by us when the conditions for reopening are...
Code of Federal Regulations, 2012 CFR
2012-04-01
...initial or reconsidered determination made by us ever appropriate? 418.1345 Section...initial or reconsidered determination made by us ever appropriate? We may reopen an initial or reconsidered determination made by us when the conditions for reopening are...
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.
Orbit determination error analysis and station-keeping for liberation point trajectories
NASA Astrophysics Data System (ADS)
Gordon, Steven Craig
In the elliptic restricted three-body problem (ER3BP), the two primary masses are assumed to be in known elliptic orbits about their common center of mass. The third (infinitesimal) mass may be positioned near one of the five known Lagrange points located in the coordinate system rotating with the primaries. The bounded motion of the infinitesimal mass relative to a Lagrange point can then be computed. In particular, for the Sun-Earth plus Moon three-body system (where the Earth plus Moon barycenter is treated as one primary mass), orbits in the vicinity of the Lagrange point L1 between the Sun and the Earth are ideal for the study of solar-terrestrial interactions. A quasi-periodic 'Lissajous' trajectory and a much larger, nearly periodic 'halo-type' orbit are used in this effort as nominal paths near L1 in the Sun-Earth plus Moon ER3BP. Trajectory determination for a spacecraft that moves under the influence of the two-body system of forces will be affected by many error sources, including tracking errors and modeling uncertainty. Orbit determination error analysis seeks to quantify the impact of these errors. Covariance analysis is a method of error analysis used in this effort to predict state vector error levels. After a predetermined tracking period, using a selected range and range-rate tracking schedule, specific covariance matrix entries are used to compute standard deviations for each of the six states. The results of error analysis using the Kalman and batch weighted least squares filters are compared, and covariance analysis is used to incorporate additional error sources, such as solar radiation pressure uncertainty. The means and the probability distributions of these state errors are tested using statistical hypothesis tests and goodness of fit tests, respectively. The state error levels are then used in Monte Carlo simulations of three station-keeping methods - two delta velocity controllers and an on/off controller developed from a state feedback algorithm. The total control cost from a single station-keeping simulation is a random variable, and a group of 30 such simulations is used as a random sample. Station-keeping costs for a spacecraft near Lissajous and halo-type nominal paths are then compared using statistical hypothesis tests.
Munoz, Douglas Perry
Abstract In previous studies of saccadic eye movement reaction time, the manipulation of initial eye position re-vealed a behavioral bias that facilitates the initiation of movements towards of positions that the eyes can take in the orbits delimits the extent of visual exploration by head
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 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 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.
20 CFR 408.1006 - What is the effect of an initial determination?
Code of Federal Regulations, 2011 CFR
2011-04-01
...determination? 408.1006 Section 408.1006 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SPECIAL BENEFITS FOR CERTAIN WORLD WAR II VETERANS Determinations and the Administrative Review Process Introduction, Definitions, and Initial...
20 CFR 408.1003 - Which administrative actions are initial determinations?
Code of Federal Regulations, 2011 CFR
2011-04-01
...determinations? 408.1003 Section 408.1003 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SPECIAL BENEFITS FOR CERTAIN WORLD WAR II VETERANS Determinations and the Administrative Review Process Introduction, Definitions, and Initial...
20 CFR 408.1006 - What is the effect of an initial determination?
Code of Federal Regulations, 2010 CFR
2010-04-01
...determination? 408.1006 Section 408.1006 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SPECIAL BENEFITS FOR CERTAIN WORLD WAR II VETERANS Determinations and the Administrative Review Process Introduction, Definitions, and Initial...
18 CFR 701.204 - Time limits for WRC initial determinations regarding requests for information.
Code of Federal Regulations, 2011 CFR
2011-04-01
...determinations regarding requests for information. 701.204 Section 701...RESOURCES COUNCIL COUNCIL ORGANIZATION Availability of Information § 701.204 Time limits...determinations regarding requests for information. (a) An initial...
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.
Precise Orbit Determination of LAGEOS satellites: results on fundamental physics and perspectives
NASA Astrophysics Data System (ADS)
Peron, Roberto; Lucchesi, David
2012-07-01
The LAGEOS satellites, launched for geodynamics and geophysics purposes, are offering also an outstanding test bench to fundamental physics. Indeed, their physical characteristics, as well as those of their orbits, and the availability of high--quality tracking data provided by the International Laser Ranging Service, allow for precise tests of gravitational theories. In this talk recent work on data analysis will be presented. A fairly large amount of LAGEOS and LAGEOS II Satellite Laser Ranging data has been analyzed with NASA/GSFC Geodyn II software, using a set of dedicated models for satellite dynamics, and the related post--fit residuals have been analyzed. In particular, general relativistic effects leave peculiar imprint on nodal longitude, argument of perigee and inclination behaviour, which have been used to obtain precise estimates of the related parameters. The most precise --- as today --- estimate of the effects on argument of perigee has been obtained, providing a direct measurement of the relativistic ``Schwarzschild'' precession in the field of the Earth. At the same time the constraints on a non--Newtonian (i.e. Yukawa--type) gravitational dynamics have been improved. The measurement error budget will be discussed, emphasizing the role of gravitational and, especially, of non--gravitational forces modeling on the overall precise orbit determination quality, as well as on future new measurements and constraints of the gravitational interaction.
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.
Designing Delta-DOR acquisition strategies to determine highly elliptical earth orbits
NASA Technical Reports Server (NTRS)
Frauenholz, R. B.
1986-01-01
Delta-DOR acquisition strategies are designed for use in determining highly elliptical earth orbits. The requirements for a possible flight demonstration are evaluated for the Charged Composition Explorer spacecraft of the Active Magnetospheric Particle Tracer Explorers. The best-performing strategy uses data spanning the view periods of two orthogonal baselines near the same orbit periapse. The rapidly changing viewing geometry yields both angular position and velocity information, but each observation may require a different reference quasar. The Delta-DOR data noise is highly dependent on acquisition geometry, varying several orders of magnitude across the baseline view periods. Strategies are selected to minimize the measurement noise predicted by a theoretical model. Although the CCE transponder is limited by S-band and a small bandwidth, the addition of Delta-DOR to coherent Doppler and range improves the one-sigma apogee position accuracy by more than an order of magnitude. Additional Delta-DOR accuracy improvements possible using dual-frequency (S/X) calibration, increased spanned bandwidth, and water-vapor radiometry are presented for comparison. With these benefits, the residual Delta-DOR data noise is primarily due to quasar position uncertainties.
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.
Orbit Determination Processes for the Navigation of the Cassini-Huygens Mission
NASA Technical Reports Server (NTRS)
Antreasian, P.G.; Ardalan, S.M.; Beswick, R.M.; Criddle, K.E.; Ionasescu, R.; Jacobson, R.A.; Jones, J.B.; MacKenzie, R.A.; Parcher, D.W.; Pelletier, F.J.; Roth, D.C.; Thompson, P.F.; Vaughan, A.T.
2008-01-01
Deep space navigation, particularly the Orbit Determination (OD) operations of Cassini at Saturn, cannot easily be automated due to the complex dynamical environment in which the spacecraft flies; however several sub-processes are automated. The Cassini OD operations are often faced with unique challenges that require more than routine procedures. The OD Team is staffed appropriately to meet the demanding schedules and allow some level of flexibility. This paper will discuss how the OD processes are developed and the seven-member OD team is scheduled to support efficient and accurate Cassini navigation operations. Also discussed will be the requirements of the radio-metric Doppler and range tracking data acquired via the Deep Space Network and the optical navigation images of the satellites to support the daily OD operations. Furthermore, the reliability of the OD solutions, which is ensured within the framework of the OD processes, will be explained.
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 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.
Solar panel orientation derived from DORIS and GPS precise orbit determination
NASA Astrophysics Data System (ADS)
Gobinddass, Marie-Line; Willis, Pascal; Haines, Bruce
Solar radiation pressure is currently a limiting factor in Precise Orbit Determination. To cope with model uncertainties, solar radiation pressure models are usually rescaled with an empirical Cr parameter. All DORIS data available from all satellites since 1993 have been reprocessed by estimating Cr coefficients on a daily basis. These time series usually show a signal depending on the satellite. Lower orbiting satellites (SPOTs and Envisat) show larger variations, especially around 11-year maximum, with a clear annual signal. Higher-altitude DORIS satellites used for altimetry (TOPEX/Poseidon, Jason-1, -2) show a better consistency. We focus here on the only 2 significant discontinuities observed in all the Cr time series: A 20% change in January 2008 for SPOT-5 and a 7% change on July, 1993 for TOPEX/Poseidon. While the SPOT-5 disconti-nuity, can be totally explained by an effective solar panel re-orientation done by CNES for this satellite (40 degrees over a few days), no similar explanation has been provided until now for TOPEX/Poseidon. Early GPS data for this satellite will be re-processed to determine whether a similar discontinuity in Cr is detected. As July 27, 1993 was the first day of solar panel re-orientation, we propose the hypothesis of a small mis-alignment of the TOPEX/Poseidon solar panel before this epoch. Consequences of ignoring these discontinuities for SPOT-5 and TOPEX/Poseidon for the DORIS geodetic results (tracking station coordinates, geocenter mo-tion) are investigated. Easy way to overcome these problems in future DORIS reprocessing activities of the International DORIS Service (IDS) are also discussed.
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.
NASA Astrophysics Data System (ADS)
Song, Young-Joo; Ahn, Sang-il; Sim, Eun-Sup
2014-09-01
In this paper, a brief but essential development strategy for the lunar orbit determination system is discussed to prepare for the future Korea's lunar missions. Prior to the discussion of this preliminary development strategy, technical models of foreign agencies for the lunar orbit determination system, tracking networks to measure the orbit, and collaborative efforts to verify system performance are reviewed in detail with a short summary of their lunar mission history. Covered foreign agencies are European Space Agency, Japan Aerospace Exploration Agency, Indian Space Research Organization and China National Space Administration. Based on the lessons from their experiences, the preliminary development strategy for Korea's future lunar orbit determination system is discussed with regard to the core technical issues of dynamic modeling, numerical integration, measurement modeling, estimation method, measurement system as well as appropriate data formatting for the interoperability among foreign agencies. Although only the preliminary development strategy has been discussed through this work, the proposed strategy will aid the Korean astronautical society while on the development phase of the future Korea's own lunar orbit determination system. Also, it is expected that further detailed system requirements or technical development strategies could be designed or established based on the current discussions.
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 [Earth-Life Science Institute (WPI-ELSI), Tokyo Institute of Technology, Ookayama, Meguro district, Tokyo 152-8551 (Japan)
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.
Schlaf, R.; Merritt, C. D.; Picciolo, L. C.; Kafafi, Z. H.
2001-08-15
We determined the orbital lineup of the tris (8-hydroxyquinolinato) gallium (Gaq{sub 3})/Mg interface using combined x-ray and ultraviolet photoemission spectroscopy (XPS and UPS) measurements. The Gaq{sub 3}/Mg system is a prototypical model structure for organic electron/low work function electrode transporting materials interfaces found in organic light emitting diodes (OLED). A Gaq{sub 3} thin film was grown in 15 steps on a previously sputter-cleaned Mg substrate starting at a 1 Aa nominal thickness up to a final thickness of 512 Aa. Before, and in between the growth steps, the sample surface was characterized by XPS and UPS. The results indicate the formation of a reaction layer of about 12 Aa thickness at the Mg interface, which resulted in a 0.96 V interface dipole potential. At Gaq{sub 3} coverages higher than 256 Aa, a strong charging shift occurred in the overlayer related UPS-emission lines, which was identified by measuring the high binding energy cutoff (secondary edge) of both the XP and UP spectra. The several magnitudes different x-ray and ultraviolet source photon intensities allow pinpointing charging shifts with high sensitivity. Due to the low work function of the reacted interface layer, the Gaq{sub 3} electronic states are aligned at a binding energy below the substrate Fermi edge that exceeds the magnitude of the optical gap between the highest occupied and lowest unoccupied molecular orbitals (HOMO and LUMO). This allowed the conclusion that the ground state exciton binding energy of Gaq{sub 3} needs to be larger than 0.43 eV. Based on these considerations, the lowest possible electron injection barrier matching the experimental data was estimated to be 0.15 eV. {copyright} 2001 American Institute of Physics.
NASA Astrophysics Data System (ADS)
Kapper, M.; Sampl, M.; Plettemeier, D.; Rucker, H. O.; Maksimovic, M.
2012-12-01
We provide updated quasi-static and high-frequency analysis and calibration of the Radio and Plasma Wave (RPW) antenna system aboard the Solar Orbiter spacecraft, whose aim is to determine the dynamics of solarwind plasma and the electric and magnetic fields in the near-Sun heliosphere with a full suite of in-situ and remote sensing instruments. The mission is planned to be launched in 2017 with a spacecraft trajectory of partial co-rotation with the Sun. Finding the true reception properties of the RPW requires not only accurate modeling of the three cylindrical antennas forming the electric sensors, but also of the spacecraft body and attached hardware. As the spacecraft will be within 0.25 AU to the sun, the antennas will also be exposed to high solar radiation and subject to strong thermal stress loads which we consider as well. The current analysis is an important step forward as our models are based on the latest spacecraft specifications which are in an advanced development stage with much of the configuration having already been finalized by ESA. The current focus is on obtaining calibration results from numerical computation of the governing field equations, resulting in surface current distributions from which the effective axes and lengths and impedance matrices for the quasi-static range are derived. This improves the instrument's performance in both remote sensing as well as in-situ capabilities. This will also complement efforts to compare with experimental data that will soon be available, including results from anechoic chamber studies performed with scale models, providing an important benchmark for the numerical results. The current calibration results are to provide useful input to goniopolarimetry techniques like polarization analysis, direction finding and ray tracing, all of which depend crucially on the effective axes, thus providing significant improvements to the corresponding data analysis.Preliminary mesh of the Solar Orbiter used for numerical calculations.
A computerized survey of a selected subset of initial orbit determination methods
Cleveland, Durand Ennis
1972-01-01
~ 1, 2. (70) The difference of the true anomalies between the two radius vectors is given by 20 1 kv = sin (~ ~-) (l-c. os hv) J 12211 (71) - f the motion is direct, and Av 1 -I 12 21 2 21 sin [-( )(1-cos 6v) (72) if the motion is retrograde... = (r r )lr 0 0 0 0 (117) and the magnitude of the velocity vector is 9 0 (118) Writing r and r in terms of the central time, f r + g r 1 o 1 o (119) r 2 2 0 2 0 (120) where [13] 2 dt. + ~2 P 1. o i ? (3uq -15 up +u) dt + 1 2 2 4 24 o o 0 0 0 i...
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.
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.
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.
NASA Astrophysics Data System (ADS)
Mangold, N.
2003-01-01
Lobate debris aprons, known to be geomorphic landform indicators of the presence of ground ice, are of special interest for future missions devoted to the research of water on Mars. Lobate debris aprons in fretted terrains of Deuteronilus and Protonilus Mensae (35°-50°N) show typical convex shapes interpreted to be the result of viscous deformation. At the scale of Mars Orbiter Camera (MOC) high-resolution images the surface of these debris aprons shows complex patterns with small pits and buttes. These patterns can be explained by the mantling of dust, the accumulation of interstitial ice, and the subsequent removal of ice by sublimation. The sublimation of the ground ice is especially initiated and accelerated by subsurface heterogeneities like fractures. Theoretical quantification of sublimation rates therefore minimizes sublimation, which is not a homogeneous process, at least over the landforms studied. Crater counts show that the sublimation occurred in the last tens of millions of years up to the recent past. In the point of view of future searching of subsurface ice, only surface layers are submitted to sublimation favoring the conservation of ground ice in deeper layers since the formation of the landform. The geophysical survey of lobate debris aprons would give interesting insights into the subsurface distribution of ice and its seasonal variations, especially in order to measure current sublimation of ground ice.
A review of averaging techniques and their application to orbit determination systems
NASA Technical Reports Server (NTRS)
Velez, C. E.; Fuchs, A. J.
1974-01-01
The theory of numerical averaging and analytical averaging will be reviewed and the application of these techniques to orbit and parameter estimation problems will be presented. Comparisons will be made between utilizing mean elements versus tracking data as the observation types. Results will be presented comparing the accuracy and efficiency of the combined orbit estimation and orbit prediction problem using averaged equations of motion, the Cowell equations of motion and the Brouwer general perturbation theory. The problem of converting the averaged element space back to osculating element space for orbit operations will also be discussed.
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.
Federal Register 2010, 2011, 2012, 2013, 2014
2012-11-19
...Review an Initial Determination Granting a Motion By Complainant To Terminate the Investigation...administrative law judge (``ALJ'') granting a motion by complainant to terminate the investigation...notice of investigation named Research In Motion Ltd. of Ontario, Canada; Research...
Determination of an Initial Mesh Density for Finite Element Computations via Data Mining
R Kanapady; S K Bathina; K K Tamma; C Kamath; V Kumar
2001-01-01
Numerical analysis software packages which employ a coarse first mesh or an inadequate initial mesh need to undergo a cumbersome and time consuming mesh refinement studies to obtain solutions with acceptable accuracy. Hence, it is critical for numerical methods such as finite element analysis to be able to determine a good initial mesh density for the subsequent finite element computations
49 CFR 7.31 - What time limits apply to DOT with respect to initial determinations?
Code of Federal Regulations, 2014 CFR
2014-10-01
... 2014-10-01 2014-10-01 false What time limits apply to DOT with respect to initial determinations...Transportation PUBLIC AVAILABILITY OF INFORMATION Time Limits § 7.31 What time limits apply to DOT with respect to initial...
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.
Grizzle, Jessy W.
A Restricted PoincarÂ´e Map for Determining Exponentially Stable Periodic Orbits in Systems the PoincarÂ´e return map around a fixed point and evaluating its eigenvalues. However, in feedback design exponentially stable--recomputing and re-linearizing the PoincarÂ´e return map at each iteration can be very
G. H. Spalding
1985-01-01
The procedures used for orbit determination and control of the AMPTE UK Satellite (UKS) are described and the operational experience is reviewed. In particular, details are given about how the separation between UKS and the companion Ion Release Module satellite was controlled to meet the scientific requirements of the AMPTE mission, utilising a cold gas thruster system on board UKS.
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.
B. Morris; J. W. Grizzle
2005-01-01
Systems with impulse effects form a special class of hybrid systems that consist of an ordinary, time-invariant differential equation (ODE), a co-dimension one switching surface, and a re-initialization rule. The exponential stability of a periodic orbit in a C1-nonlinear systems with impulse effects can be studied by linearizing the Poincaré return map around a fixed point and evaluating its eigenvalues.
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.
Federal Register 2010, 2011, 2012, 2013, 2014
2012-02-15
...337-TA-755] Certain Starter Motors and Alternators; Determination...the investigation as to respondent Electric Motor Service, Inc. (EMS) of Logan...after importation of certain starter motors and alternators that by reason...
12 CFR 709.8 - Administrative appeal of the initial determination.
Code of Federal Regulations, 2011 CFR
2011-01-01
...administrative appeal shall rest solely with the Board, which...initial determination shall rest solely upon the claimant...determination to which the claimant objects and the alleged error in such...requests, shall apply with equal force to any such amendment or...
12 CFR 709.8 - Administrative appeal of the initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
...administrative appeal shall rest solely with the Board, which...initial determination shall rest solely upon the claimant...determination to which the claimant objects and the alleged error in such...requests, shall apply with equal force to any such amendment or...
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.
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%.
Initial on-orbit performance analysis of Inverted Metamorphic (IMM3J) solar cells on MISSE-7
Kenneth M. Edmondson; Alex Howard; Paul Hausgen; Phillip Jenkins; Dhananjay Bhusari; Sonya Wierman; Shoghig Mesropian; Daniel C. Law; Rina Bardfield; Richard R. King; Nasser H. Karam
2011-01-01
Prototype Inverted Metamorphic (IMM3J) cells were grown and fabricated and assembled onto an experimental flight coupon for inclusion on the 7th Materials International Space Station Experiment (MISSE-7). This paper examines the first eleven months of on-orbit data of prototype IMM3J solar cells in a low earth orbit (LEO) environment. The prototype IMM3J solar cells show excellent electrical and mechanical stability
Code of Federal Regulations, 2014 CFR
2014-04-01
...initially determine and then adjust expected levels of performance for the core performance...initially determine and then adjust expected levels of performance for the core performance...will undertake to agree upon expected levels of performance for each core...
Code of Federal Regulations, 2012 CFR
2012-04-01
...initially determine and then adjust expected levels of performance for the core performance...initially determine and then adjust expected levels of performance for the core performance...will undertake to agree upon expected levels of performance for each core...
Code of Federal Regulations, 2013 CFR
2013-04-01
...initially determine and then adjust expected levels of performance for the core performance...initially determine and then adjust expected levels of performance for the core performance...will undertake to agree upon expected levels of performance for each core...
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.
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 Technologies’ 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.
12 CFR 563b.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2010 CFR
2010-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 563b...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 192.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2012 CFR
2012-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 192...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 563b.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2011 CFR
2011-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 563b...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 192.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2014 CFR
2014-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 192...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 563b.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2012 CFR
2012-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 563b...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 563b.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2014 CFR
2014-01-01
...true How do I determine the initial balances of liquidation sub-accounts? 563b...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 192.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2013 CFR
2013-01-01
...false How do I determine the initial balances of liquidation sub-accounts? 192...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
12 CFR 563b.460 - How do I determine the initial balances of liquidation sub-accounts?
Code of Federal Regulations, 2013 CFR
2013-01-01
...true How do I determine the initial balances of liquidation sub-accounts? 563b...460 How do I determine the initial balances of liquidation sub-accounts? ...You determine the initial sub-account balance for a savings account held by an...
Determination of spin-orbit coupling contributions in the framework of density functional theory.
Chiodo, Sandro; Russo, Nino
2008-04-30
We present a noniterative method to calculate spin-orbit coupling by means of a theoretical approach that provides the use of the full Breit-Pauli operator. This method was applied to compute one and two-electron spin-orbit coupling contributions between singlet and triplet, and doublet and doublet states, respectively. These states have been represented by monodeterminantal wave functions and optimized using the PW91 gradient-corrected exchange-correlation functional and the hybrid B3LYP one. They have been supplied by the conventional density functional theory packages, and thus coupled by our spin-orbit coupling code. Different size basis sets have been employed and the obtained results have been compared with the corresponding ones provided by some of the already existing methods and with the experimental data. They have been found to be in good quantitative agreement. PMID:17963223
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.
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.
First assessment of GPS-based reduced dynamic orbit determination on TOPEX/Poseidon
NASA Technical Reports Server (NTRS)
Yunck, T. P.; Bertiger, W. I.; Wu, S. C.; Bar-Sever, Y. E.; Christensen, E. J.; Haines, B. J.; Lichten, S. M.; Muellerschoen, R. J.; Vigue, Y.; Willis, P.
1994-01-01
The reduced dynamic Global Positioning System (GPS) tracking technique has been applied for the first time as part of the GPS experiment on TOPEX/Poseidon. This technique employs local geometric position corrections to reduce orbit errors caused by the mismodeling of satellite forces. Results for a 29-day interval in early 1993 are evaluated through postfit residuals and formal errors, comparison with GPS and laser/DORIS dynamic solutions, comparisons on 6-hr overlaps of adjacent 30-hr data arcs, altimetry closure and crossover analysis. Reduced dynamic orbits yield slightly better crossover agreement than other techniques and appear to be accurate in altitude to about 3 cm RMS.
NASA Technical Reports Server (NTRS)
Vonbun, F. O.; Argentiero, P. D.; Schmid, P. E.
1978-01-01
The results of the ATS-6/GEOS-3 and the ATS-6/NIMBUS-6 satellite-to-satellite tracking orbit determination experiments are reported. The tracking systems used in these experiments differ from the Tracking and Data Relay Satellite System (TDRSS), primarily in the use of one rather than two synchronous relay satellites. However, the simulations mentioned indicate that the insights gained from the experiments with regard to proper data reduction techniques and expected results are applicable to the TDRSS.
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.
Andysz, Aleksandra; Najder, Anna; Merecz-Kot, Dorota
2014-01-01
Appropriate distribution of time and energy between work and personal life poses a challenge to many working people. Unfortunately, many professionally active people experience work-family conflict. In order to minimize it, employees are offered various solutions aimed at reconciling professional and private spheres (work-life balance (WLB) initiatives). The authors attempt to answer what makes employees use WLB initiatives and what influences the decision to reject the available options. The review is based on the articles published after 2000, searched by Google Scholar and Web of Knowledge with use of the key words: work-life balance, work-family conflict, work-life balance initiatives, work-life balance initiatives use, use of WLB solutions. We focused on organizational and individual determinants of WLB initiatives use, such as organizational culture, stereotypes and values prevailing in the work environment that may result in stigmatization of workers - flexibility stigma. We discuss the reasons why supervisors and co-workers stigmatize their colleagues, and what are the consequences of experiencing such stigmatization. Among the individual determinants of WLB initiatives use, we have inter alia focused on the preference for integration vs. separation of the spheres of life. The presented material shows that social factors - cultural norms prevailing in a society, relationships in the workplace and individual factors, such as the level of self-control - are of equal importance for decisions of using WLB initiatives as their existence. Our conclusion is that little attention has been paid to the research on determinants of WLB initiatives use, especially to individual ones. PMID:24834699
Predisposing, enabling and pregnancy-related determinants of late initiation of prenatal care.
Beeckman, Katrien; Louckx, Fred; Putman, Koen
2011-10-01
Prenatal care is important for the health and wellbeing of women and their babies. There is international consensus that prenatal care should begin in the first trimester. This study aims to analyze the effects of predisposing, enabling and pregnancy-related determinants of late prenatal care initiation. In this prospective observational study, 333 women were recruited consecutively at the beginning of their prenatal care trajectory. Data was collected on the timing of the first prenatal visit and on socio-demographic and pregnancy-related characteristics, using a semi-structured interview. A multivariate binominal logistic regression was applied to analyze independent effects on late initiation of prenatal care. Bivariately late initiation of care was associated with being inactive on the labor market, non-European origin, not having lived in Belgium since birth, low income, receiving welfare benefits, not having a regular obstetrician and experiencing difficulties getting a first appointment. When adjusting for all determinants, our multivariate analyses showed that late initiation was associated with non-European origin, low income and not having a regular obstetrician. This study shows that late initiation of prenatal care is associated with predisposing and enabling determinants. In order to ensure timely initiation of care, policy-makers should focus on encouraging women to have a regular prenatal care provider before pregnancy and taking steps in lowering out-of-pocket fees for low-income women. Future research is needed to examine whether these determinants are associated with initiation of care only or whether they play a role in the pregnancy follow-up as well. PMID:20661634
Social determinants of health in Canada: Are healthy living initiatives there yet? A policy analysis
2012-01-01
Introduction Preventative strategies that focus on addressing the social determinants of health to improve healthy eating and physical activity have become an important strategy in British Columbia and Ontario for combating chronic diseases. What has not yet been examined is the extent to which healthy living initiatives implemented under these new policy frameworks successfully engage with and change the social determinants of health. Methods Initiatives active between January 1, 2006 and September 1, 2011 were found using provincial policy documents, web searches, health organization and government websites, and databases of initiatives that attempted to influence to nutrition and physical activity in order to prevent chronic diseases or improve overall health. Initiatives were reviewed, analyzed and grouped using the descriptive codes: lifestyle-based, environment-based or structure-based. Initiatives were also classified according to the mechanism by which they were administered: as direct programs (e.g. directly delivered), blueprints (or frameworks to tailor developed programs), and building blocks (resources to develop programs). Results 60 initiatives were identified in Ontario and 61 were identified in British Columbia. In British Columbia, 11.5% of initiatives were structure-based. In Ontario, of 60 provincial initiatives identified, 15% were structure-based. Ontario had a higher proportion of direct interventions than British Columbia for all intervention types. However, in both provinces, as the intervention became more upstream and attempted to target the social determinants of health more directly, the level of direct support for the intervention lessened. Conclusions The paucity of initiatives in British Columbia and Ontario that address healthy eating and active living through action on the social determinants of health is problematic. In the context of Canada's increasingly neoliberal political and economic policy, the public health sector may face significant barriers to addressing upstream determinants in a meaningful way. If public health cannot directly affect broader societal conditions, interventions should be focused around advocacy and education about the social determinants of health. It is necessary that health be seen for what it is: a political matter. As such, the health sector needs to take a more political approach in finding solutions for health inequities. PMID:22889402
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.
Calibration and validation of individual GOCE accelerometers by precise orbit determination
NASA Astrophysics Data System (ADS)
Visser, P. N. A. M.; IJssel, J. A. A. van den
2015-09-01
The European Space Agency Gravity field and steady-state Ocean Circular Explorer (GOCE) carries a gradiometer consisting of three pairs of accelerometers in an orthogonal triad. Precise GOCE science orbit solutions (PSO), which are based on satellite-to-satellite tracking observations by the Global Positioning System and which are claimed to be at the few cm precision level, can be used to calibrate and validate the observations taken by the accelerometers. This has been done for each individual accelerometer by a dynamic orbit fit of the time series of position co-ordinates from the PSOs, where the accelerometer observations represent the non-gravitational accelerations. Since the accelerometers do not coincide with the center of mass of the GOCE satellite, the observations have to be corrected for rotational and gravity gradient terms. This is not required when using the so-called common-mode accelerometer observations, provided the center of the gradiometer coincides with the GOCE center of mass. Dynamic orbit fits based on these common-mode accelerations therefore served as reference. It is shown that for all individual accelerometers, similar dynamic orbit fits can be obtained provided the above-mentioned corrections are made. In addition, accelerometer bias estimates are obtained that are consistent with offsets in the gravity gradients that are derived from the GOCE gradiometer observations.
The orbit of asteroid (99942) Apophis as determined from optical and radar observations
T. A. Vinogradova; O. M. Kochetova; Yu. A. Chernetenko; V. A. Shor; E. I. Yagudina
2008-01-01
The results of improving the orbit accuracy for the asteroid Apophis and the circumstances of its approach to Earth in 2029 are described. Gravitational perturbations from all of the major planets and Pluto, Ceres, Pallas, and Vesta are taken into account in the equations of motion of the asteroid. Relativistic perturbations from the Sun and perturbations due to the oblateness
B. Ts. Bakhshiyan; A. A. Sukhanov; K. S. Fedyaev
2010-01-01
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.
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.
K. H. Ilk; A. Löcher
There is a need for a proper validation procedure of gravity field solutions, especially for those high precise ones which are derived from the dedicated gravity field missions as CHAMP, GRACE and in future GOCE. In this paper the balance equations for the energy and for the energy exchange of a satellite orbiting around the Earth are proposed as analysis
The Synergy of Direct Imaging and Astrometry for Orbit Determination of exo-Earths
Shao, Michael; Pan, Xiaopei
2010-01-01
The holy grail of exoplanet searches is an exo-Earth, an Earth mass planet in the habitable zone around a nearby star. Mass is the most important parameter of a planet and can only be measured by observing the motion of the star around the planet-star center of mass. A single image of a planet, however, does not provide evidence that the planet is Earth mass or that it is in a habitable zone orbit. The planet's orbit, however, can be measured either by imaging the planet at multiple epochs or by measuring the position of the star at multiple epochs by space-based astrometry. The measurement of an exo-planet's orbit by direct imaging is complicated by a number of factors: (1) the inner working angle (IWA); (2) the apparent brightness of the planet depending on the orbital phase; (3) confusion arising from the presence of multiple planets; and (4) the planet-star contrast. In this paper we address the question: "Can a prior astrometric mission that can identify which stars have Earthlike planets significantly i...
D. Giovannini; F. M. Miatto; J. Romero; S. M. Barnett; J. P. Woerdman; M. J. Padgett
2012-07-10
The Shannon dimensionality of orbital angular momentum (OAM) entanglement produced in spontaneous parametric down-conversion can be probed by using multi-sector phase analysers. We demonstrate a spatial light modulator-based implementation of these analysers, and use it to measure a Schmidt number of about 50.
NASA Astrophysics Data System (ADS)
Baur, O.; Wirnsberger, H.; Kirchner, G.; Schreiber, K. U.; Riede, W.
2014-12-01
In the last decades, Satellite Laser Ranging (SLR) has been proven to be one of the most valuable techniques to improve our understanding of the Earth's shape, gravity field and rotational motion. Apart from the ''traditional'' tasks of SLR, in the recent past the portfolio has been considerably widened. Especially the SLR-based orbit determination and orbit prediction of space debris objects is gaining increasing attention. The reliable and accurate orbit determination/prediction of debris objects is of crucial importance for any effort towards debris collision avoidance. Hitherto, this task is performed by space surveillance networks using ground-based radar tracking and passive optical tracking with telescopes. These methods, however, are characterized by low accuracy, ranging from several hundreds of meters to a few kilometers for low-Earth orbiting objects. As demonstrated successfully, SLR has the potential to considerably improve these numbers, and hence to improve debris surveillance. One of the most severe limitations of the technique is the sparse network of SLR stations, and hence the sparseness of tracking data. Against this background, we propose to extend the existing SLR network by passive telescopes in combination with multi-static observations. Multi-static observations means that an object is tracked by only one active SLR station, but the diffusely reflected photons are detected at several passive stations (in case of only one passive station the measurements are referred to as bi-static). In order to demonstrate the performance of the principle, we analyzed (i) two-way laser ranges from Graz (3 passes), and (ii) two-way laser ranges from Graz (3 passes) in combination with bi-static measurements between the SLR stations Graz and Wettzell (3 passes); our investigations focus on the defunct ENVISAT satellite. We found the orbit determination/prediction results including bi-static observations to outperform single-station results by about one order of magnitude. For scenario (ii), our orbit prediction error ranges from 40m (prediction age 1 day) to 450m (prediction age 7 days). From our first experiments we conclude that bi-/multi-static SLR is capable to meet the requirements for debris surveillance, but might also be beneficial for more ''traditional'' applications.
19 CFR 210.66 - Initial determination concerning temporary relief; Commission action thereon.
Code of Federal Regulations, 2011 CFR
2011-04-01
...day is a Saturday, Sunday, or Federal holiday, the initial determination must...intermediary Saturdays, Sundays, and Federal holidays shall be included. If the last day of...period is a Saturday, Sunday, or Federal holiday as defined in § 201.14(a) of...
Precision orbit determination using the Tracking and Data Relay Satellite System (TDRSS)
NASA Technical Reports Server (NTRS)
Roesset, P.; Lundberg, J. B.; Tapley, B. D.
1993-01-01
The growth of the Tracking and Data Relay Satellite System (TDRSS) is the result of a greater reliance on the systems to provide nearly global coverage for relaying data from environmental satellites and to reduce or eliminate the reliance on global networks of tracking ground stations. Tracking data collected by TDRSS is often used to compute orbital solutions for moperational mission requirements. Investigations are in progress that seek to assess the feasibility of extending the use of tracking data collected by TDRSS as a means for computing precise orbital solutions. Specifically, this investigation will use covariance analysis techniques to evaluate this extended capability as applied to the TOPEX/Poseidon mission. This study will complement other investigations which carry out similar assessments of TDRSS using actual tracking data. This paper presents some preliminary results for Cycle 5 of the TOPEX/Poseidon mission using simulated two-way range-rate measurements.
B. Ts. Bakhshiyan; A. A. Sukhanov; K. S. Fedyaev
2010-01-01
An analysis of the existing astrometric and radar observations of the Apophis asteroid is performed. On the basis of this\\u000a analysis, characteristics of future measurements of the asteroid orbit and limitation on their conduction are accepted. A\\u000a proposed launching of a spacecraft to the asteroid in order to obtain high-accuracy measurements of its distance and radial\\u000a velocity is also considered.
The orbit of asteroid (99942) Apophis as determined from optical and radar observations
T. A. Vinogradova; O. M. Kochetova; Yu. A. Chernetenko; V. A. Shor; E. I. Yagudina
2008-01-01
The results of improving the orbit accuracy for the asteroid Apophis and the circumstances of its approach to Earth in 2029\\u000a are described. Gravitational perturbations from all of the major planets and Pluto, Ceres, Pallas, and Vesta are taken into\\u000a account in the equations of motion of the asteroid. Relativistic perturbations from the Sun and perturbations due to the oblateness
J-Adaptive estimation with estimated noise statistics. [for orbit determination
NASA Technical Reports Server (NTRS)
Jazwinski, A. H.; Hipkins, C.
1975-01-01
The J-Adaptive estimator described by Jazwinski and Hipkins (1972) is extended to include the simultaneous estimation of the statistics of the unmodeled system accelerations. With the aid of simulations it is demonstrated that the J-Adaptive estimator with estimated noise statistics can automatically estimate satellite orbits to an accuracy comparable with the data noise levels, when excellent, continuous tracking coverage is available. Such tracking coverage will be available from satellite-to-satellite tracking.
Factorization methods for precision satellite orbit determination foi HC\\/MF
G. J. Bierman; L. A. Campbell; W. A. Feess
1981-01-01
State-of-the-art square root information filtering and smoothing technology that is incorporated into the Aerospace TRACE orbital analysis program are presented. Topics include a pseudo-epoch state batch-sequential filter formulation, techniques for inclusion of Markov process noise models, a variable dimension filter structure that accomodates state vectors of large size, and inclusion of a GPS second-order Markov clock model within the framework
High-precision onboard orbit determination for small satellites - the GPS-based XNSon X-SAT
NASA Astrophysics Data System (ADS)
Gill, E.; Montenbruck, O.; Arichandran, K.; Tan, S.H.; Bretschneider
2004-11-01
X-SAT is a mini-satellite developed by the Satellite Engineering Centre of the Nanyang Technological University at Singapore. The focus of the technology- driven mission is the high-resolution remote sensing of the Southeast Asian region for environmental monitoring. To achieve the ambitious mission objectives, the GPS-based X-SAT Navigation System (XNS) will provide high-precision onboard orbit determination solutions as well as orbit forecasts. With a targeted real-time position accuracy of about 1-2 m 3D r.m.s., the XNS provides an unprecedented accuracy performance and thus enables the support of any satellite mission which requires precise onboard position knowledge.
AN ANALYTIC METHOD TO DETERMINE HABITABLE ZONES FOR S-TYPE PLANETARY ORBITS IN BINARY STAR SYSTEMS
Eggl, Siegfried; Pilat-Lohinger, Elke; Gyergyovits, Markus; Funk, Barbara; Georgakarakos, Nikolaos E-mail: elke.pilat-lohinger@univie.ac.at
2012-06-10
With more and more extrasolar planets discovered in and around binary star systems, questions concerning the determination of the classical habitable zone have arisen. Do the radiative and gravitational perturbations of the second star influence the extent of the habitable zone significantly, or is it sufficient to consider the host star only? In this article, we investigate the implications of stellar companions with different spectral types on the insolation a terrestrial planet receives orbiting a Sun-like primary. We present time-independent analytical estimates and compare them to insolation statistics gained via high precision numerical orbit calculations. Results suggest a strong dependence of permanent habitability on the binary's eccentricity, as well as a possible extension of habitable zones toward the secondary in close binary systems.
The Synergy of Direct Imaging and Astrometry for Orbit Determination of Exo-Earths
NASA Astrophysics Data System (ADS)
Shao, Michael; Catanzarite, Joseph; Pan, Xiaopei
2010-09-01
The holy grail of exoplanet searches is an exo-Earth, an Earth mass planet in the habitable zone (HZ) around a nearby star. Mass is one of the most important characteristics of a planet and can only be measured by observing the motion of the star around the planet-star center of gravity. The planet's orbit can be measured either by imaging the planet at multiple epochs or by measuring the position of the star at multiple epochs by space-based astrometry. The measurement of an exoplanet's orbit by direct imaging is complicated by a number of factors. One is the inner working angle (IWA). A space coronagraph or interferometer imaging an exo-Earth can separate the light from the planet from the light from the star only when the star-planet separation is larger than the IWA. Second, the apparent brightness of a planet depends on the orbital phase. A single image of a planet cannot tell us whether the planet is in the HZ or distinguish whether it is an exo-Earth or a Neptune-mass planet. Third is the confusion that may arise from the presence of multiple planets. With two images of a multiple planet system, it is not possible to assign a dot to a planet based only on the photometry and color of the planet. Finally, the planet-star contrast must exceed a certain minimum value in order for the planet to be detected. The planet may be unobservable even when it is outside the IWA, such as when the bright side of the planet is facing away from us in a "crescent" phase. In this paper we address the question: "Can a prior astrometric mission that can identify which stars have Earth-like planets significantly improve the science yield of a mission to image exo-Earths?" In the case of the Occulting Ozone Observatory, a small external occulter mission that cannot measure spectra, we find that the occulter mission could confirm the orbits of ~4 to ~5 times as many exo-Earths if an astrometric mission preceded it to identify which stars had such planets. In the case of an internal coronagraph we find that a survey of the nearest ~60 stars could be done with a telescope half the size if an astrometric mission had first identified the presence of Earth-like planets in the HZ and measured their orbital parameters.
NASA Technical Reports Server (NTRS)
Kirschner, S. M.; Beri, A. C.; Broaddus, S. R.; Doll, C. E.
1990-01-01
In order to validate the operational and computational capabilities of the Preliminary Orbit Determination System (PODS), tests were performed using tracking measurements for several systems including the ERB satellite, the SMM, the STS and Landsat-4. POD procedures are utilized to generate a state vector following an unplanned orbital perturbation or spacecraft maneuver, when an estimation process such as a differential correction orbit determination cannot obtain a solution. Results are presented to demonstrate POD for several situations involving different qualities of a priori target state vectors, data type combinations, data arc lengths, and mixtures of single-TDRS, dual-TDRS, and GSTDN measurements. The system's ability to determine accurately the state vector for the spacecraft and the effectiveness of the solution screening process are discussed. It is shown that PODS is capable of determining a spacecraft vector when differential correction orbit determination processes fail.
Determination of laser damage initiation probability and growth on fused silica scratches
Norton, M A; Carr, C W; Cross, D A; Negres, R A; Bude, J D; Steele, W A; Monticelli, M V; Suratwala, T I
2010-10-26
Current methods for the manufacture of optical components inevitably leaves a variety of sub-surface imperfections including scratches of varying lengths and widths on even the finest finishes. It has recently been determined that these finishing imperfections are responsible for the majority of laser-induced damage for fluences typically used in ICF class lasers. We have developed methods of engineering subscale parts with a distribution of scratches mimicking those found on full scale fused silica parts. This much higher density of scratches provides a platform to measure low damage initiation probabilities sufficient to describe damage on large scale optics. In this work, damage probability per unit scratch length was characterized as a function of initial scratch width and post fabrication processing including acid-based etch mitigation processes. The susceptibility of damage initiation density along scratches was found to be strongly affected by the post etching material removal and initial scratch width. We have developed an automated processing procedure to document the damage initiations per width and per length of theses scratches. We show here how these tools can be employed to provide predictions of the performance of full size optics in laser systems operating at 351 nm. In addition we use these tools to measure the growth rate of a damage site initiated along a scratch and compare this to the growth measured on an isolated damage site.
NASA Technical Reports Server (NTRS)
Pepper, Stephen V.
2011-01-01
The destruction rates of a perfluoropolyether (PFPE) lubricant, Krytox 143AC, subjected to rolling contact with 440C steel in a spiral orbit tribometer at room temperature have been evaluated as a function of test environment. The rates in ultrahigh vacuum, 0.213 kPa (1.6 torr) oxygen and one atmosphere of dry nitrogen were about the same. Water vapor in the test environment-a few ppm in one atmosphere of nitrogen-reduced the destruction rate by up to an order of magnitude. A similar effect of water vapor was found for the destruction rate of Pennzane 2001A, an unformulated multiply alkylated cyclopentane (MAC) hydrocarbon oil.
The dynamics of global positioning system orbits and the determination of precise ephemerides
NASA Technical Reports Server (NTRS)
Colombo, Oscar L.
1989-01-01
The suggestion made on the basis of the analytical orbit perturbation theory that the errors in the ephemerides of the GPS satellites are due mostly to resonant effects that can be corrected by adjusting a few parameters in a empirical acceleration formula is tested using simulations and actual data analysis. Data from the Spring 1985 Experiment were used to calculate improved ephemerides, and these ephemerides were used in the estimation of the coordinates of GPS stations within the continental United States, previously positioned with VLBI. The results of this test support the idea that the errors are mostly of a resonant nature and can be corrected.
NASA Technical Reports Server (NTRS)
Vonbraun, C.; Reigber, Christoph
1994-01-01
In the spring of 1993, the MOMS-02 (modular Optoelectronic Multispectral Scanner) camera, as part of the second German Spacelab mission aboard STS-55, successfully took digital threefold stereo images of the surface of the Earth. While the mission is experimental in nature, its primary goals are to produce high quality maps and three-dimensional digital terrain models of the Earth's surface. Considerable improvement in the quality of the terrain model can be attained if information about the position and attitude of the camera is included during the adjustment of the image data. One of the primary sources of error in the Shuttle's position is due to the significant attitude maneuvers conducted during the course of the mission. Various arcs, using actual Tracking and Data Relay Satellite (TDRSS) Doppler data of STS-55, were processed to determine how effectively empirical force modeling could be used to solve for the radial, transverse, and normal components of the orbit perturbations caused by these routine maneuvers. Results are presented in terms of overlap-orbit differences in the three components. Comparisons of these differences, before and after the maneuvers are estimated, show that the quality of an orbit can be greatly enhanced with this technique, even if several maneuvers are present. Finally, a discussion is made of some of the difficulties encountered with this approach, and some ideas for future studies are presented.
Determination of an Initial Mesh Density for Finite Element Computations via Data Mining
Kanapady, R; Bathina, S K; Tamma, K K; Kamath, C; Kumar, V
2001-07-23
Numerical analysis software packages which employ a coarse first mesh or an inadequate initial mesh need to undergo a cumbersome and time consuming mesh refinement studies to obtain solutions with acceptable accuracy. Hence, it is critical for numerical methods such as finite element analysis to be able to determine a good initial mesh density for the subsequent finite element computations or as an input to a subsequent adaptive mesh generator. This paper explores the use of data mining techniques for obtaining an initial approximate finite element density that avoids significant trial and error to start finite element computations. As an illustration of proof of concept, a square plate which is simply supported at its edges and is subjected to a concentrated load is employed for the test case. Although simplistic, the present study provides insight into addressing the above considerations.
Using Data to Determine the Initial Conditions in Heavy Ion Collisions
NASA Astrophysics Data System (ADS)
Soltz, Ron; Garishvili, Irakli; Abelev, Betty
2012-10-01
We have developed a framework, the Comprehensive Heavy Ion Model Evaluation Reporting Algorithm (CHIMERA) to determine the optimal model and initial conditions of heavy ion collisions by comparing to data from a variety of observables. We have used this framework to study simple participant and binary collisions scaling in the presence of pre-equilibrium flow in the context of the VH2 2D+1 viscous hydrodynamic model with UrQMD afterburner for data from RHIC. We have also used this framework to explore the significance of variations in the equation of state. We have recently begun to apply this framework to a new hydro-solver tools known as CHOMBO, which incorporates adaptive mesh refinement techniques that are well suited to the study of initial state fluctuations. We will review results from using CHIMERA with VH2, and discuss future plans for using CHOMBO to study initial state fluctuations.
Code of Federal Regulations, 2011 CFR
2011-04-01
...initial determination that is based on a more recent tax year? 418.1230 Section...Adjustment Amount Determinations Using A More Recent Tax Year's Modified Adjusted Gross...initial determination that is based on a more recent tax year? (a) When you...
Code of Federal Regulations, 2011 CFR
2011-04-01
...amount initial determination based on a more recent tax year? 418.2230 Section...Coverage Premiums Determinations Using A More Recent Tax Year's Modified Adjusted Gross...amount initial determination based on a more recent tax year? We will follow...
Code of Federal Regulations, 2010 CFR
2010-04-01
...initial determination that is based on a more recent tax year? 418.1230 Section...Adjustment Amount Determinations Using A More Recent Tax Year's Modified Adjusted Gross...initial determination that is based on a more recent tax year? (a) When you...
NASA Technical Reports Server (NTRS)
Luthcke, S. B.; Marshall, J. A.
1992-01-01
The TOPEX/Poseidon spacecraft was launched on August 10, 1992 to study the Earth's oceans. To achieve maximum benefit from the altimetric data it is to collect, mission requirements dictate that TOPEX/Poseidon's orbit must be computed at an unprecedented level of accuracy. To reach our pre-launch radial orbit accuracy goals, the mismodeling of the radiative nonconservative forces of solar radiation, Earth albedo an infrared re-radiation, and spacecraft thermal imbalances cannot produce in combination more than a 6 cm rms error over a 10 day period. Similarly, the 10-day drag modeling error cannot exceed 3 cm rms. In order to satisfy these requirements, a 'box-wing' representation of the satellite has been developed in which, the satellite is modelled as the combination of flat plates arranged in the shape of a box and a connected solar array. The radiative/thermal nonconservative forces acting on each of the eight surfaces are computed independently, yielding vector accelerations which are summed to compute the total aggregate effect on the satellite center-of-mass. Select parameters associated with the flat plates are adjusted to obtain a better representation of the satellite acceleration history. This study analyzes the estimation of these parameters from simulated TOPEX/Poseidon laser data in the presence of both nonconservative and gravity model errors. A 'best choice' of estimated parameters is derived and the ability to meet mission requirements with the 'box-wing' model evaluated.
Determination of dynamic fracture-initiation toughness using a novel impact bend test procedure
T. Yokoyama
1993-01-01
A novel impact bend test procedure is described for determining the dynamic fracture-initiation toughness, K[sub Id], at a loading rate (stress intensity factor rate), K[sub I], of the order of 10[sup 6] MPa [radical]m\\/s. A special arrangement of the split Hopkinson pressure bar is adopted to measure accurately dynamic loads applied to a fatigue-precracked bend specimen. The dynamic stress intensity
Code of Federal Regulations, 2010 CFR
2010-04-01
... Customs Duties UNITED STATES INTERNATIONAL TRADE COMMISSION INVESTIGATIONS OF UNFAIR PRACTICES IN IMPORT TRADE ADJUDICATION AND ENFORCEMENT...initial determination raises a policy matter which the Commission...
Orbital geometry determined by orthogonal high-order harmonic polarization components
Hijano, Eliot; Serrat, Carles; Gibson, George N.; Biegert, Jens
2010-04-15
We study the polarization state of high-order harmonics produced by linearly polarized light interacting with two-center molecules. By generating high-harmonic 'polarization maps' from Radon transformations of excited electronic wave functions, we show that the polarization of the harmonic radiation can be linked to the geometry of the molecular orbital. While in the Radon transformation the plane-wave approximation for the rescattered electron is implicitly assumed, numerical solutions of the two-dimensional time-dependent Schro{center_dot}{center_dot}dinger equation, in which this approximation is not made, confirm the validity of this topological connection. We also find that measuring two orthogonal amplitude components of the harmonics provides a method for quantum tomography that substantially improves the quality of reconstructed molecular states.
Determining Spent Nuclear Fuel's Plutonium Content, Initial Enrichment, Burnup, and Cooling Time
Cheatham, Jesse R; Francis, Matthew W
2011-01-01
The Next Generation of Safeguards Initiative is examining nondestructive assay techniques to determine the total plutonium content in spent nuclear fuel. The goal of this research was to develop new techniques that can independently verify the plutonium content in a spent fuel assembly without relying on an operator's declarations. Fundamentally this analysis sought to answer the following questions: (1) do spent fuel assemblies contain unique, identifiable isotopic characteristics as a function of their burnup, cooling time, and initial enrichment; (2) how much variation can be seen in spent fuel isotopics from similar and dissimilar reactor power operations; and (3) what isotopes (if any) could be used to determine burnup, cooling time, and initial enrichment? To answer these questions, 96,000 ORIGEN cases were run that simulated typical two-cycle operations with burnups ranging from 21,900 to 72,000 MWd/MTU, cooling times from 5 to 25 years, and initial enrichments between 3.5 and 5.0 weight percent. A relative error coefficient was determined to show how numerically close a reference solution has to be to another solution for the two results to be indistinguishable. By looking at the indistinguishable solutions, it can be shown how a precise measurement of spent fuel isotopics can be inconclusive when used in the absence of an operator's declarations. Using this Method of Indistinguishable Solutions (MIS), we evaluated a prominent method of nondestructive analysis - gamma spectroscopy. From this analysis, a new approach is proposed that demonstrates great independent forensic examination potential for spent nuclear fuel by examining both the neutron emissions of Cm-244 and the gamma emissions of Cs-134 and Eu-154.
Determination of the initial beam parameters in Monte Carlo linac simulation.
Aljarrah, Khaled; Sharp, Greg C; Neicu, Toni; Jiang, Steve B
2006-04-01
For Monte Carlo linac simulations and patient dose calculations, it is important to accurately determine the phase space parameters of the initial electron beam incident on the target. These parameters, such as mean energy and radial intensity distribution, have traditionally been determined by matching the calculated dose distributions with the measured dose distributions through a trial and error process. This process is very time consuming and requires a lot of Monte Carlo simulation experience and computational resources. In this paper, we propose an easy, efficient, and accurate method for the determination of the initial beam parameters. We hypothesize that (1) for one type of linacs, the geometry and material of major components of the treatment head are the same; the only difference is the phase space parameters of the initial electron beam incident on the target, and (2) most linacs belong to a limited number of linac types. For each type of linacs, Monte Carlo treatment planning system (MC-TPS) vendors simulate the treatment head and calculate the three-dimensional (3D) dose distribution in water phantom for a grid of initial beam energies and radii. The simulation results (phase space files and dose distribution files) are then stored in a data library. When a MC-TPS user tries to model their linac which belongs to the same type, a standard set of measured dose data is submitted and compared with the calculated dose distributions to determine the optimal combination of initial beam energy and radius. We have applied this method to the 6 MV beam of a Varian 21EX linac. The linac was simulated using EGSNRC/BEAM code and the dose in water phantom was calculated using EGSNRC/DOSXYZ. We have also studied issues related to the proposed method. Several common cost functions were tested for comparing measured and calculated dose distributions, including chi2, mean absolute error, dose difference at the penumbra edge point, slope of the dose difference of the lateral profile, and the newly proposed Kappaalpha factor (defined as the fraction of the voxels with absolute dose difference less than alpha%). It was found that the use of the slope of the lateral profile difference or the difference of the penumbra edge points may lead to inaccurate determination of the initial beam parameters. We also found that in general the cost function value is very sensitive to the simulation statistical uncertainty, and there is a tradeoff between uncertainty and specificity. Due to the existence of statistical uncertainty in simulated dose distributions, it is practically impossible to determine the best energy/radius combination; we have to accept a group of energy/radius combinations. We have also investigated the minimum required data set for accurate determination of the initial beam parameters. We found that the percent depth dose curves along or only a lateral profile at certain depth for a large field size is not sufficient and the minimum data set should include several lateral profiles at various depths as well as the central axis percent depth dose curve for a large field size. PMID:16696460
Sekanina, Zdenek; Chodas, Paul W. E-mail: Paul.W.Chodas@jpl.nasa.gov
2012-10-01
We describe the physical and orbital properties of C/2011 W3. After surviving perihelion passage, the comet was observed to undergo major physical changes. The permanent loss of the nuclear condensation and the formation of a narrow spine tail were observed first at Malargue, Argentina, on December 20 and then systematically at Siding Spring, Australia. The process of disintegration culminated with a terminal fragmentation event on December 17.6 UT. The postperihelion dust tail, observed for {approx}3 months, was the product of activity over <2 days. The nucleus' breakup and crumbling were probably caused by thermal stress due to the penetration of the intense heat pulse deep into the nucleus' interior after perihelion. The same mechanism may be responsible for cascading fragmentation of sungrazers at large heliocentric distances. The delayed response to the hostile environment in the solar corona is at odds with the rubble-pile model, since the residual mass of the nucleus, estimated at {approx}10{sup 12} g (equivalent to a sphere 150-200 m across) just before the terminal event, still possessed nontrivial cohesive strength. The high production rates of atomic oxygen, observed shortly after perihelion, are compatible with a subkilometer-sized nucleus. The spine tail-the product of the terminal fragmentation-was a synchronic feature, whose brightest part contained submillimeter-sized dust grains, released at velocities of up to 30 m s{sup -1}. The loss of the nuclear condensation prevented an accurate orbital-period determination by traditional techniques. Since the missing nucleus must have been located on the synchrone, whose orientation and sunward tip have been measured, we compute the astrometric positions of this missing nucleus as the coordinates of the points of intersection of the spine tail's axis with the lines of forced orbital-period variation, derived from the orbital solutions based on high-quality preperihelion astrometry from the ground. The resulting orbit gives 698 {+-} 2 yr for the osculating orbital period, showing that C/2011 W3 is the first member of the expected new, 21st-century cluster of bright Kreutz-system sungrazers, whose existence was predicted by these authors in 2007. From the spine tail's evolution, we determine that its measured tip, populated by dust particles 1-2 mm in diameter, receded antisunward from the computed position of the missing nucleus. The bizarre appearance of the comet's dust tail in images taken only hours after perihelion with the coronagraphs on board the SOHO and STEREO spacecraft is readily understood. The disconnection of the comet's head from the tail released before perihelion and an apparent activity attenuation near perihelion have a common cause-sublimation of all dust at heliocentric distances smaller than about 1.8 solar radii. The tail's brightness is strongly affected by forward scattering of sunlight by dust. From an initially broad range of particle sizes, the grains that were imaged the longest had a radiation-pressure parameter {beta} {approx_equal} 0.6, diagnostic of submicron-sized silicate grains and consistent with the existence of the dust-free zone around the Sun. The role and place of C/2011 W3 in the hierarchy of the Kreutz system and its genealogy via a 14th-century parent suggest that it is indirectly related to the celebrated sungrazer X/1106 C1, which, just as the first-generation parent of C/2011 W3, split from a common predecessor during the previous return to perihelion.
NASA Astrophysics Data System (ADS)
Sekanina, Zdenek; Chodas, Paul W.
2012-10-01
We describe the physical and orbital properties of C/2011 W3. After surviving perihelion passage, the comet was observed to undergo major physical changes. The permanent loss of the nuclear condensation and the formation of a narrow spine tail were observed first at Malargue, Argentina, on December 20 and then systematically at Siding Spring, Australia. The process of disintegration culminated with a terminal fragmentation event on December 17.6 UT. The postperihelion dust tail, observed for ~3 months, was the product of activity over <2 days. The nucleus' breakup and crumbling were probably caused by thermal stress due to the penetration of the intense heat pulse deep into the nucleus' interior after perihelion. The same mechanism may be responsible for cascading fragmentation of sungrazers at large heliocentric distances. The delayed response to the hostile environment in the solar corona is at odds with the rubble-pile model, since the residual mass of the nucleus, estimated at ~1012 g (equivalent to a sphere 150-200 m across) just before the terminal event, still possessed nontrivial cohesive strength. The high production rates of atomic oxygen, observed shortly after perihelion, are compatible with a subkilometer-sized nucleus. The spine tail—the product of the terminal fragmentation—was a synchronic feature, whose brightest part contained submillimeter-sized dust grains, released at velocities of up to 30 m s-1. The loss of the nuclear condensation prevented an accurate orbital-period determination by traditional techniques. Since the missing nucleus must have been located on the synchrone, whose orientation and sunward tip have been measured, we compute the astrometric positions of this missing nucleus as the coordinates of the points of intersection of the spine tail's axis with the lines of forced orbital-period variation, derived from the orbital solutions based on high-quality preperihelion astrometry from the ground. The resulting orbit gives 698 ± 2 yr for the osculating orbital period, showing that C/2011 W3 is the first member of the expected new, 21st-century cluster of bright Kreutz-system sungrazers, whose existence was predicted by these authors in 2007. From the spine tail's evolution, we determine that its measured tip, populated by dust particles 1-2 mm in diameter, receded antisunward from the computed position of the missing nucleus. The bizarre appearance of the comet's dust tail in images taken only hours after perihelion with the coronagraphs on board the SOHO and STEREO spacecraft is readily understood. The disconnection of the comet's head from the tail released before perihelion and an apparent activity attenuation near perihelion have a common cause—sublimation of all dust at heliocentric distances smaller than about 1.8 solar radii. The tail's brightness is strongly affected by forward scattering of sunlight by dust. From an initially broad range of particle sizes, the grains that were imaged the longest had a radiation-pressure parameter ? ~= 0.6, diagnostic of submicron-sized silicate grains and consistent with the existence of the dust-free zone around the Sun. The role and place of C/2011 W3 in the hierarchy of the Kreutz system and its genealogy via a 14th-century parent suggest that it is indirectly related to the celebrated sungrazer X/1106 C1, which, just as the first-generation parent of C/2011 W3, split from a common predecessor during the previous return to perihelion.
Host and bacterial determinants of initial severity and outcome of Escherichia coli sepsis.
Jauréguy, F; Carbonnelle, E; Bonacorsi, S; Clec'h, C; Casassus, P; Bingen, E; Picard, B; Nassif, X; Lortholary, O
2007-09-01
A 1-year prospective cohort study of all episodes of Escherichia coli bacteraemia in two French university hospitals was conducted to assess simultaneously the influence of host and bacterial determinants on the initial severity and outcome of E. coli sepsis. Clinical data (community-acquired/nosocomial infection, immune status, underlying disease, primary source of infection, severity sepsis scoring and outcome), phylogenetic groups (A, B1, D and B2), nine virulence factors (VFs) (papC, papGII, papGIII, sfa/foc, hlyC, cnf1, iucC, fyuA and iroN) and the antibiotic susceptibility of isolates were investigated. All VFs except iucC were significantly more prevalent (p <0.05) among the B2 group isolates. The non-B2 isolates were more frequently resistant to antibiotics than were B2 isolates (p <0.05). There were significantly more B2 isolates from immunocompetent than from immunocompromised patients (p <0.05). No bacterial or host determinants influenced the initial severity of sepsis. Multivariate analysis revealed that the presence of papGIII, septic shock at baseline and a non-urinary tract origin of sepsis were associated independently with a fatal outcome (p 0.04, <0.0001 and 0.04, respectively). A factorial analysis of correspondence allowed two populations of isolates to be distinguished: those belonging to the B2 group were associated more frequently with susceptibility to antibiotics, community-acquired infection, a urinary tract origin and immunocompetent hosts; those belonging to the A, B1 or D groups were associated more frequently with resistance to antibiotics, a nosocomial origin, a non-urinary tract source and immunocompromised hosts. Although no influence of host or bacterial determinants on the initial severity of sepsis was detected, bacterial and host determinants both influenced the outcome of E. coli sepsis significantly. PMID:17617183
Kim, Julianne; Iaboni, Dolores C.; Walker, Scott E.; Elligsen, Marion; Dunn, Michael S.; Allen, Vanessa G.; Simor, Andrew
2014-01-01
Variability in neonatal vancomycin pharmacokinetics and the lack of consensus for optimal trough concentrations in neonatal intensive care units pose challenges to dosing vancomycin in neonates. Our objective was to determine vancomycin pharmacokinetics in neonates and evaluate dosing regimens to identify whether practical initial recommendations that targeted trough concentrations most commonly used in neonatal intensive care units could be determined. Fifty neonates who received vancomycin with at least one set of steady-state levels were evaluated retrospectively. Mean pharmacokinetic values were determined using first-order pharmacokinetic equations, and Monte Carlo simulation was used to evaluate initial dosing recommendations for target trough concentrations of 15 to 20 mg/liter, 5 to 20 mg/liter, and ?20 mg/liter. Monte Carlo simulation revealed that dosing by mg/kg of body weight was optimal where intermittent dosing of 9 to 12 mg/kg intravenously (i.v.) every 8 h (q8h) had the highest probability of attaining a target trough concentration of 15 to 20 mg/liter. However, continuous infusion with a loading dose of 10 mg/kg followed by 25 to 30 mg/kg per day infused over 24 h had the best overall probability of target attainment. Initial intermittent dosing of 9 to 15 mg/kg i.v. q12h was optimal for target trough concentrations of 5 to 20 mg/liter and ?20 mg/liter. In conclusion, we determined that the practical initial vancomycin dose of 10 mg/kg vancomycin i.v. q12h was optimal for vancomycin trough concentrations of either 5 to 20 mg/liter or ?20 mg/liter and that the same initial dose q8h was optimal for target trough concentrations of 15 to 20 mg/liter. However, due to large interpatient vancomycin pharmacokinetic variability in neonates, monitoring of serum concentrations is recommended when trough concentrations between 15 and 20 mg/liter or 5 and 20 mg/liter are desired. PMID:24614381
Determination of dynamic fracture-initiation toughness using a novel impact bend test procedure
Yokoyama, T. (Osaka Univ. (Japan). Faculty of Engineering Okayama Univ. of Science (Japan). Dept. of Mechanical Engineering)
1993-11-01
A novel impact bend test procedure is described for determining the dynamic fracture-initiation toughness, K[sub Id], at a loading rate (stress intensity factor rate), K[sub I], of the order of 10[sup 6] MPa [radical]m/s. A special arrangement of the split Hopkinson pressure bar is adopted to measure accurately dynamic loads applied to a fatigue-precracked bend specimen. The dynamic stress intensity factor history for the bend specimen is evaluated by means of a dynamic finite element technique. The onset of crack initiation is detected using a string gage attached on the side of the specimen near a crack tip. The value of K[sub Id] is determined from the critical dynamic stress intensity factor at crack initiation. A series of dynamic fracture tests is carried out on a 7075-T6 aluminum alloy, a Ti-6246 alloy and an AISI 4340 steel. The K[sub Id] values obtained for the three structural materials are compared with the corresponding values obtained under quasi-static loading conditions.
Conservative force model performance for TOPEX/Poseidon precision orbit determination
NASA Technical Reports Server (NTRS)
Marshall, J. Andrew; Luthcke, Scott B.
1993-01-01
The TOPEX/Poseidon spacecraft was launched on August 10, 1992 to study the Earth's oceans. To achieve maximum benefit from the altimetric data collected, mission requirements dictate that TOPEX/Poseidon's orbit must be computed at an unprecedented level of accuracy. In order to satisfy these requirements, a model which accounts for the satellite's complex geometry, attitude, and surface properties has been developed. This `box-wing' representation treats the spacecraft as the combination of flat plates arranged in the shape of a box and a connecetd solar array. The nonconservative forces acting on each of the eight surfaces are computed independently, yielding vector accelerations which are summed to compute the total aggregate effect on the satellite center-of-mass. Parameters associated with each flat plate were derived from a finite element analysis of the spacecraft. Certain parameters can be inferred from tracking data and have been adjusted to obtain a better representation of the satellite acceleration history. Changes in the nominal mission profile and the presence of an `anomalistic' force have complicated this tuning process. Model performance, parameter sensitivities, and the `anomalistic' force will be discussed.
GIOVE-B solar radiation pressure modeling for precise orbit determination
NASA Astrophysics Data System (ADS)
Steigenberger, Peter; Montenbruck, Oliver; Hugentobler, Urs
2015-03-01
Previous studies have identified systematic errors in the orbit and clock estimates of the GIOVE and Galileo IOV satellites in the order of ± 20 cm. These errors are visible as periodic variations in the Satellite Laser Ranging (SLR) and clock residuals. For IOV, these variations could be attributed to the contribution of a stretched satellite body and it was shown that a simple a priori box model for the solar radiation pressure can significantly reduce these errors. GIOVE-B has similar dimensions as the IOV satellites but its orientation is different: for GIOVE-B the narrow side of the satellite points to the Earth rather than the longitudinal side. In addition, an extra plate carrying, amongst others, the laser retro reflector array is mounted on the spacecraft introducing shadowing effects. These features are considered with a simple box-plate model. This model reduces the periodic clock errors and the SLR residual RMS of GIOVE-B by a factor of two. Most importantly, the box-plate model reduces the SLR offset from 11 cm to less than 1 cm. The largest part of this reduction comes from considering the plate and its shadowing effects.
Orbital angular momentum in electron diffraction and its use to determine chiral crystal symmetries
Juchtmans, Roeland
2015-01-01
In this work we present an alternative way to look at electron diffraction in a transmission electron microscope. In stead 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 mulsti-slice simulations on $\\alpha$-q...
Conversion of Osculating Orbital Elements to Mean Orbital Elements
NASA Technical Reports Server (NTRS)
Der, Gim J.; Danchick, Roy
1996-01-01
Orbit determination and ephemeris generation or prediction over relatively long elapsed times can be accomplished with mean elements. The most simple and efficient method for orbit determination, which is also known as epoch point conversion, performs the conversion of osculating elements to mean elements by iterative procedures. Previous epoch point conversion methods are restricted to shorter elapsed times with linear convergence. The new method presented in this paper calculates an analytic initial guess of the unknown mean elements from a first order theory of secular perturbations and computes a transition matrix with accurate numerical partials. It thereby eliminates the problem of an inaccurate initial guess and an identity transition matrix employed by previous methods. With a good initial guess of the unknown mean elements and an accurate transition matrix, converging osculating elements to mean elements can be accomplished over long elapsed times with quadratic convergence.
Code of Federal Regulations, 2014 CFR
2014-04-01
...Is reopening of an initial or reconsidered determination made by us ever appropriate? 418.2345 Section 418.2345 Employees...Is reopening of an initial or reconsidered determination made by us ever appropriate? We will follow the rules in §...
Code of Federal Regulations, 2011 CFR
2011-04-01
...Is reopening of an initial or reconsidered determination made by us ever appropriate? 418.2345 Section 418.2345 Employees...Is reopening of an initial or reconsidered determination made by us ever appropriate? We will follow the rules in §...
Code of Federal Regulations, 2012 CFR
2012-04-01
...Is reopening of an initial or reconsidered determination made by us ever appropriate? 418.2345 Section 418.2345 Employees...Is reopening of an initial or reconsidered determination made by us ever appropriate? We will follow the rules in §...
Code of Federal Regulations, 2013 CFR
2013-04-01
...Is reopening of an initial or reconsidered determination made by us ever appropriate? 418.2345 Section 418.2345 Employees...Is reopening of an initial or reconsidered determination made by us ever appropriate? We will follow the rules in §...
Jin Wen; Masahiro Yamamoto; Ting Wei
2012-01-01
In this article, we consider an inverse heat conduction problem of determining a time-dependent unknown heat source and unknown initial temperature by means of observations of the temperature at the final time and temperature profile at one fixed point over the time interval. We prove that the heat source and initial temperature can be determined uniquely from two kinds of
NASA Technical Reports Server (NTRS)
Lemoine, F. G.; Rowlands, D. D.; Luthcke, S. B.; Zelensky, N. P.; Chinn, D. S.; Pavlis, D. E.; Marr, G. C.
2001-01-01
The U.S. Navy's GEOSAT Follow-On Spacecraft was launched on February 10, 1998 and the primary objective of the mission was 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. The spacecraft is tracked by satellite laser ranging (SLR) and Doppler (Tranet-style) beacons. Although a limited amount of GPS data were obtained, the primary mode of tracking remains satellite laser ranging. In this paper, we report on progress in orbit determination for GFO using GFO/GFO and TOPEX/GFO altimeter crossovers. We have tuned the nonconservative force model for GFO and the gravity model using SLR, Doppler and altimeter crossover data spanning over one year. Preliminary results show that the predicted radial orbit error from the gravity field covariance to 70x70 on GEOSAT was reduced from 2.6 cm in EGM96 to 1.9 cm with the addition of only five months of the GFO SLR and GFO/GFO crossover data. Further progress is possible with the addition of more data, particularly the TOPEX/GFO crossovers. We will evaluate the tuned GFO gravity model (a derivative of EGM96) using altimeter data from the GEOSAT mission. In January 2000, a limited quantity of GPS data were obtained. We will use these GPS data in conjunction with the SLR and altimeter crossover data obtained over the same time span to compute quasi-reduced dynamic orbits which will also aid in the evaluation of the tuned GFO geopotential model.
NASA Technical Reports Server (NTRS)
Rutledge, Sharon K.
1999-01-01
Spacecraft in low Earth orbit (LEO) are subjected to many components of the environment, which can cause them to degrade much more rapidly than intended and greatly shorten their functional life. The atomic oxygen, ultraviolet radiation, and cross contamination present in LEO can affect sensitive surfaces such as thermal control paints, multilayer insulation, solar array surfaces, and optical surfaces. The LEO Spacecraft Materials Test (LEO-SMT) program is being conducted to assess the effects of simulated LEO exposure on current spacecraft materials to increase understanding of LEO degradation processes as well as to enable the prediction of in-space performance and durability. Using ground-based simulation facilities to test the durability of materials currently flying in LEO will allow researchers to compare the degradation evidenced in the ground-based facilities with that evidenced on orbit. This will allow refinement of ground laboratory test systems and the development of algorithms to predict the durability and performance of new materials in LEO from ground test results. Accurate predictions based on ground tests could reduce development costs and increase reliability. The wide variety of national and international materials being tested represent materials being functionally used on spacecraft in LEO. The more varied the types of materials tested, the greater the probability that researchers will develop and validate predictive models for spacecraft long-term performance and durability. Organizations that are currently participating in the program are ITT Research Institute (USA), Lockheed Martin (USA), MAP (France), SOREQ Nuclear Research Center (Israel), TNO Institute of Applied Physics (The Netherlands), and UBE Industries, Ltd. (Japan). These represent some of the major suppliers of thermal control and sensor materials currently flying in LEO. The participants provide materials that are exposed to selected levels of atomic oxygen, vacuum ultraviolet radiation, contamination, or synergistic combined environments at the NASA Lewis Research Center. Changes in characteristics that could affect mission performance or lifetime are then measured. These characteristics include changes in mass, solar absorptance, and thermal emittance. The durability of spacecraft materials from U.S. suppliers is then compared with those of materials from other participating countries. Lewis will develop and validate performance and durability prediction models using this ground data and available space data. NASA welcomes the opportunity to consider additional international participants in this program, which should greatly aid future spacecraft designers as they select materials for LEO missions.
The Determination and Long Term Integration of the Orbits of Caliban and Sycorax
NASA Astrophysics Data System (ADS)
Jacobson, R. A.
1999-09-01
The first 2 irregular satellites of Uranus, Caliban and Sycorax, were discovered in late 1997 (Gladman et al. 1998 Nature 392, 897). Subsequently, pre-discovery observations of both satellites were found on plates taken by D. Cruikshank in June of 1984. Recently, P. Nicholson, D. Tholen, and W. Offutt provided observations which they made in late 1998 at Palomar Mountain, Mauna Kea, and Cloudcroft, respectively. I fit a numerical integration perturbed by the Sun, Jupiter, Saturn, and Neptune to the set of available observations. For the 47 observations of Caliban the respective rms values of the Delta alpha cos delta and Delta delta residuals are 0\\farcs60 and 0\\farcs32, and for the 103 observations of Sycorax the analogous values are 0\\farcs57 and 0\\farcs59. I extended the integration to span a 6000 year period and computed osculating orbital elements at yearly intervals. The table below contains the mean values of the elements over the 6000 years, the sidereal period, and the precession periods of the argument of periapsis and longitude of the ascending node. The osculating elements (except for a) exhibit a significant long period oscillation with a period roughly half that of the argument of periapsis. Element & Caliban Sycorax a (km) & 7166840 & 12191450 e & 0.168 & 0.520 omega (deg) & 153.32 & 20.99 i (deg) & 140.93 & 156.99 Omega (deg) & 168.00 & 263.69 P (day) & 579.46 & 1283.26 P_? (yr) & 8900 & 1390 P_? (yr) & 6700 & 1860 Ephemerides for the satellites are available electronically from the JPL Horizons on-line solar system data and ephemeris computation service.
NASA Technical Reports Server (NTRS)
Mohammed, Priscilla N.; Piepmeier, Jeffrey R.; Johnson, Joel T.; Aksoy, Mustafa; Bringer, Alexandra
2015-01-01
The Soil Moisture Active Passive (SMAP) mission, launched in January 2015, provides global measurements of soil moisture using a microwave radiometer. SMAPs radiometer passband lies within the passive frequency allocation. However, both unauthorized in-band transmitters as well as out-of-band emissions from transmitters operating at frequencies adjacent to this allocated spectrum have been documented as sources of radio frequency interference (RFI) to the L-band radiometers on SMOS and Aquarius. The spectral environment consists of high RFI levels as well as significant occurrences of low level RFI equivalent to 0.1 to 10 K. The SMAP ground processor reports the antenna temperature both before and after RFI mitigation is applied. The difference between these quantities represents the detected RFI level. The presentation will review the SMAP RFI detection and mitigation procedure and discuss early on-orbit RFI measurements from the SMAP radiometer. Assessments of global RFI properties and source types will be provided, as well as the implications of these results for SMAP soil moisture measurements.
Lior M. Burko
2006-04-06
Comparing the corrections to Kepler's law with orbital evolution under a self force, we extract the finite, already regularized part of the latter in a specific gauge. We apply this method to a quasi-circular orbit around a Schwarzschild black hole of an extreme mass ratio binary, and determine the first- and second-order conservative gravitational self force in a post Newtonian expansion. We use these results in the construction of the gravitational waveform, and revisit the question of the relative contribution of the self force and spin-orbit coupling.
A study to explore the use of orbital remote sensing to determine native arid plant distribution
NASA Technical Reports Server (NTRS)
Mcginnies, W. G. (principal investigator); Haase, E. F.; Musick, H. B. (compiler)
1973-01-01
The author has identified the following significant results. A theory has been developed of a method for determining the reflectivities of natural areas from ERTS-1 data. This method requires the following measurements: (1) ground truth reflectivity data from two different calibration areas; (2) radiance data from ERTS-1 MSS imagery for the same two calibration areas; and (3) radiance data from ERTS-1 MSS imagery for the area(s) in which reflectivity is to be determined. The method takes into account sun angle effects and atmospheric effects on the radiance seen by the space sensor. If certain assumptions are made, the ground truth data collection need not be simultaneous with the ERTS-1 overflight. The method allows the calculation of a conversion factor for converting ERTS-1 MSS radiance measurements of a given overflight to reflectivity values. This conversion factor can be used to determine the reflectivity of any area in the general vicinity of the calibration areas which has a relatively similar overlying atmosphere. This method, or some modification of it, may be useful in ERTS investigations which require the determination of spectral signatures of areas from spacecraft data.
Techniques for the determination of mass properties of earth-to-orbit transportation systems
NASA Technical Reports Server (NTRS)
Macconochie, I. O.; Klich, P. J.
1978-01-01
One estimating technique involves trending whereby projections of overall mass properties of vehicles are determined with few inputs. The second technique involves trending of individual subsystems using equations of the form KXN to the nth power or KX. Some constants and exponentials are provided for sample subsystems. Mass properties are reported in a format recommended by mil spec - 38310.
NASA Astrophysics Data System (ADS)
Davidhazy, Andrew
1991-04-01
The stress testing of latex condoms by an air burst procedure has been slow in gaining industry acceptance because questions have been raised regarding the influence of the test apparatus on the likelihood of breakage occurring where the condom is attached to the inflation device. It was desired to locate the areas at which the condoms tend to burst and thus corroborate or disprove these claims. Several factors associated with the bursting condom demanded the use of special instrumentation to detect arid study the burst initiation process. Microsecond duration electronic flashes were used for the initial stages of the investigation. Although the absolute point of initiation of a given burst could not be photographed, these high speed studies tend to indicate that the most likely place for high quality condoms to break is not where they are attached to the inflation device but at an intermediate area between the base and the tip of the condom. In addition, tear propagation characteristics and velocities were determined with a delayed-flash technique, a double-slit strip method and a rotating drum framing camera.
Uyttenhove, W.; Van Den Eynde, G.; Baeten, P.; Kochetkov, A.; Vittiglio, G.; Wagemans, J. [SCKCEN, Belgian Nuclear Research Centre, Boeretang 200, BE-2400 Mol (Belgium); Lathouwers, D.; Kloosterman, J. L.; Van Der Hagen, T. J. H. H.; Wols, F. [Delft Univ. of Technology, Mekelweg 15, NL-2629 JB Delft (Netherlands); Billebaud, A.; Chabod, S.; Thybault, H. E. [LPSC-CNRS-IN2P3/UJF/INPG, 53, Avenue des Martyrs, 38026 Grenoble Cedex (France); Lecouey, J. L.; Ban, G.; Lecolley, F. R.; Marie, N.; Steckmeyer, J. C. [LPC Caen, ENSICAEN/Unicersit de Caen/CNRS-IN2P3, Caen (France); Dessagne, P.; Kerveno, M. [IPHC-DRS/UdS/CNRS-IN2P3, Strasbourg (France); Mellier, F. [CEA/DEN/DER/SPEX/LPE, Cadarache 13108 Saint-Paul-les-Durance (France)
2012-07-01
Within the GUINEVERE project (Generation of Uninterrupted Intense Neutrons at the lead Venus Reactor) carried out at SCK-CEN in Mol, the continuous deuteron accelerator GENEPI-3C was coupled to the VENUS-F fast simulated lead-cooled reactor. Today the FREYA project (Fast Reactor Experiments for hYbrid Applications) is ongoing to study the neutronic behavior of this Accelerator Driven System (ADS) during different phases of operation. In particular the set-up of a monitoring system for the subcriticality of an ADS is envisaged to guarantee safe operation of the installation. The methodology for subcriticality monitoring in ADS takes into account the determination of the initial subcriticality level, the monitoring of reactivity variations, and interim cross-checking. At start-up, the Pulsed Neutron Source (PNS) technique is envisaged to determine the initial subcriticality level. Thanks to its reference critical state, the PNS technique can be validated on the VENUS-F core. A detector positioning methodology for the PNS technique is set up in this paper for the subcritical VENUS-F core, based on the reduction of higher harmonics in a static evaluation of the Sjoestrand area method. A first case study is provided on the VENUS-F core. This method can be generalised in order to create general rules for detector positions and types for full-scale ADS. (authors)
Determining the Separation and Position Angles of Orbiting Binary Stars: Comparison of Three Methods
NASA Astrophysics Data System (ADS)
Walsh, Ryan; Boule, Cory; Andrews, Katelyn; Penfield, Andrew; Ross, Ian; Lucas, Gaylon; Braught, Trisha; Harfenist, Steven; Goodale, Keith
2015-07-01
To initiate a long term binary star research program, undergraduate students compared the accuracy and ease of measuring the separations and position angles of three long period binary pairs using three different measurement techniques. It was found that digital image capture using BackyardEOS software and subsequent analysis in Adobe Photoshop was the most accurate and easiest to use of our three methods. The systems WDS J17419+7209 (STF 2241AB), WDS 19418+5032 (STFA 46AB), and WDS 16362+5255 (STF 2087AB) were found to have separations and position angles of: 30", 16°; 39.7", 133°; and 3.1", 104°, respectively. This method produced separation values within 1.3" and position angle values within 1.3° of the most recently observed values found in the Washington Double Star Catalog.
NASA Technical Reports Server (NTRS)
Boltz, F. W.
1984-01-01
An algorithm is presented for efficient p-iterative solution of the Lambert/Gauss orbit-determination problem using second-order Newton iteration. The algorithm is based on a universal transformation of Kepler's time-of-flight equation and approximate inverse solutions of this equation for short-way and long-way flight paths. The approximate solutions provide both good starting values for iteration and simplified computation of the second-order term in the iteration formula. Numerical results are presented which indicate that in many cases of practical significance (except those having collinear position vectors) the algorithm produces at least eight significant digits of accuracy with just two or three steps of iteration.
NASA Technical Reports Server (NTRS)
Todorovic-Juchniewicz, Bozenna; Sitarski, Grzegorz
1992-01-01
To improve the orbits, all the positional observations of the comets were collected. The observations were selected and weighted according to objective mathematical criteria and the mean residuals a priori were calculated for both comets. We took into account nongravitational effects in the comets' motion using Marsden's method applied in two ways: either determining the three constant parameters, A(sub 1), A(sub 2), A(sub 3) or the four parameters A, eta, I, phi connected with the rotating nucleus of the comet. To link successfully all the observations, we had to assume for both comets that A(t) = A(sub O)exp(-B x t) where B was an additional nongravitational parameter.
NASA Astrophysics Data System (ADS)
Królikowska, Ma?gorzata; Dybczy?ski, Piotr A.
2013-10-01
Dynamics of a complete sample of small perihelion distance near-parabolic comets discovered in the years 2006-2010 are studied (i.e. of 22 comets of qosc < 3.1 au). First, osculating orbits are obtained after a very careful positional data inspection and processing, including where appropriate, the method of data partitioning for determination of pre- and post-perihelion orbit for tracking then its dynamical evolution. The non-gravitational acceleration in the motion is detected for 50 per cent of investigated comets, in a few cases for the first time. Different sets of non-gravitational parameters are determined from pre- and post-perihelion data for some of them. The influence of the positional data structure on the possibility of the detection of non-gravitational effects and the overall precision of orbit determination is widely discussed. Secondly, both original and future orbits were derived by means of numerical integration of swarms of virtual comets obtained using a Monte Carlo cloning method. This method allows us to follow the uncertainties of orbital elements at each step of dynamical evolution. The complete statistics of original and future orbits that includes significantly different uncertainties of 1/a-values is presented, also in the light of our results obtained earlier. Basing on 108 comets examined by us so far, we conclude that only one of them, C/2007 W1 Boattini, seems to be a serious candidate for an interstellar comet. We also found that 53 per cent of 108 near-parabolic comets escaping in the future from the Solar system, and the number of comets leaving the Solar system as so called Oort spike comets (i.e. comets suffering very small planetary perturbations) is 14 per cent. A new method for cometary orbit quality assessment is also proposed by means of modifying the original method, introduced by Marsden, Sekanina & Everhart. This new method leads to a better diversification of orbit quality classes for contemporary comets.
NASA Astrophysics Data System (ADS)
May, C. L.; Smith Pryor, B.; Lisle, T. E.
2012-12-01
Reservoir releases on large regulated rivers are increasingly being used to rejuvenate riverine habitat downstream of dams. Determining the effective flow level is complicated by the trade-off between mobilizing bed particles and retaining coarse sediment in rivers with low sediment supply. This study determined mobilization and transport distance of bed particles using motion-sensing radio transmitting particles that approximated the reach-average D84 grain size. The distribution of shear stress at initial motion varied substantially between flood events, and suggests that the sequence of flood events and the history of underthreshold flows may be an important determinant of bed strength and thus particle mobility. In addition, particle activity was greatest on the rising limb of each flood and was maximized at near bank-full flow. Travel distances did not vary between floods when scaled by transport event duration, and a negative exponential distribution was a good fit to the data. Results of this study provide important insight into individual particle movement, which can be used to inform flow releases and understand the effects of flood magnitude on particle mobility and transport.