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.
Dealing with uncertainties in angles-only initial orbit determination
NASA Astrophysics Data System (ADS)
Armellin, Roberto; Di Lizia, Pierluigi; Zanetti, Renato
2016-08-01
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 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.
Dealing with uncertainties in angles-only initial orbit determination
NASA Astrophysics Data System (ADS)
Armellin, Roberto; Di Lizia, Pierluigi; Zanetti, Renato
2016-05-01
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 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.
Genetic Algorithm for Initial Orbit Determination with Too Short Arc
NASA Astrophysics Data System (ADS)
Li, X. R.; Wang, X.
2016-01-01
The sky surveys of space objects have obtained a huge quantity of too-short-arc (TSA) observation data. However, the classical method of initial orbit determination (IOD) can hardly get reasonable results for the TSAs. The IOD is reduced to a two-stage hierarchical optimization problem containing three variables for each stage. Using the genetic algorithm, a new method of the IOD for TSAs is established, through the selection of optimizing variables as well as the corresponding genetic operator for specific problems. Numerical experiments based on the real measurements show that the method can provide valid initial values for the follow-up work.
NASA Astrophysics Data System (ADS)
Weisman, R. M.; Majji, M.; Alfriend, K. T.
2014-02-01
This paper presents an approach to characterize the uncertainty associated with the state vector obtained from the Herrick-Gibbs orbit determination approach using transformation of variables. The approach is applied to estimate the state vector and its probability density function for objects in low Earth orbit using sparse observations. The state vector and associated uncertainty estimates are computed in Cartesian coordinates and Keplerian elements. The approach is then extended to accommodate the J_2 perturbation where the state vector is written in terms of mean orbital elements. The results obtained from the analytical approach presented in this paper are validated using Monte Carlo simulations and compared with the often utilized similarity transformation for Kepler, mean, and nonsingular elements. The measurement uncertainty characterization obtained is used to initialize conventional nonlinear filters as well as operate a Bayesian approach for orbit determination and object tracking.
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.
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
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
Short-Arc Correlation and Initial Orbit Determination For Space-Based Observations
NASA Astrophysics Data System (ADS)
Fujimoto, K.; Scheeres, D.
2011-09-01
Initial orbit determination (IOD) of space debris is an important segment of space situational awareness and is often coupled with the problem of track correlation, since in order to determine the orbit of an observed object, multiple observations must be combined. It is generally uncertain, however, whether two arbitrary tracks are of the same object. Recently, Fujimoto and Scheeres have proposed a novel and rigorous track correlation and IOD technique where each observation is assigned an “admissible region” in state space based on some physical constraints. The relationship of two observations is then determined by finding whether these regions intersect via Bayes’ rule. In this paper, we propose a new application of this method to space-based observations. Preliminary results show robustness to classically singular geometries, such as GEO-on-GEO observations. Admissible regions were first proposed by Milani et al. for heliocentric orbits, and Tommei et al. expanded this concept to Earth orbiting objects. Maruskin et al. was first to introduce the concept of intersecting multiple admissible regions to correlate tracks and obtain an initial orbit estimate, albeit the correlation was conducted in 2-dimensional subspaces of the state space. Fujimoto and Scheeres fully developed ways of characterizing intersections of admissible regions in the full 6-dimensional state space. They showed through topological arguments that a positive correlation also simultaneously provides an initial orbit estimate. A method of linearly mapping admissible regions to the state space was introduced in order to improve computational turn-around, and was validated with a series of numerical tests. For space-based observations, the observation location vector, previously assumed to be Earth-fixed, is now allowed to propagate under two-body dynamics. Several technical challenges arise when we make this change. First, the admissible region spans over a larger region in the state space
Initial results of precise orbit and clock determination for COMPASS navigation satellite system
NASA Astrophysics Data System (ADS)
Zhao, Qile; Guo, Jing; Li, Min; Qu, Lizhong; Hu, Zhigang; Shi, Chuang; Liu, Jingnan
2013-05-01
The development of the COMPASS satellite system is introduced, and the regional tracking network and data availability are described. The precise orbit determination strategy of COMPASS satellites is presented. Data of June 2012 are processed. The obtained orbits are evaluated by analysis of post-fit residuals, orbit overlap comparison and SLR (satellite laser ranging) validation. The RMS (root mean square) values of post-fit residuals for one month's data are smaller than 2.0 cm for ionosphere-free phase measurements and 2.6 m for ionosphere-free code observations. The 48-h orbit overlap comparison shows that the RMS values of differences in the radial component are much smaller than 10 cm and those of the cross-track component are smaller than 20 cm. The SLR validation shows that the overall RMS of observed minus computed residuals is 68.5 cm for G01 and 10.8 cm for I03. The static and kinematic PPP solutions are produced to further evaluate the accuracy of COMPASS orbit and clock products. The static daily COMPASS PPP solutions achieve an accuracy of better than 1 cm in horizontal and 3 cm in vertical. The accuracy of the COMPASS kinematic PPP solutions is within 1-2 cm in the horizontal and 4-7 cm in the vertical. In addition, we find that the COMPASS kinematic solutions are generally better than the GPS ones for the selected location. Furthermore, the COMPASS/GPS combinations significantly improve the accuracy of GPS only PPP solutions. The RMS values are basically smaller than 1 cm in the horizontal components and 3-4 cm in the vertical component.
NASA Astrophysics Data System (ADS)
Kelecy, Tom; Shoemaker, Michael; Jah, Moriba
2013-08-01
A break-up in Low Earth Orbit (LEO) is simulated for 10 objects having area-to-mass ratios (AMR's) ranging from 0.1-10.0 m2/kg. The Constrained Admissible Region Multiple Hypothesis Filter (CAR-MHF) is applied to determining and characterizing the orbit and atmospheric drag parameters (CdA/m) simultaneously for each of the 10 objects with no a priori orbit or drag information. The results indicate that CAR-MHF shows promise for accurate, unambiguous and autonomous determination of the orbit and drag states.
NASA Technical Reports Server (NTRS)
Jordan, J. F.; Boggs, D. H.; Born, G. H.; Christensen, E. J.; Ferrari, A. J.; Green, D. W.; Hylkema, R. K.; Mohan, S. N.; Reinbold, S. J.; Sievers, G. L.
1973-01-01
A historic account of the activities of the Satellite OD Group during the MM'71 mission is given along with an assessment of the accuracy of the determined orbit of the Mariner 9 spacecraft. Preflight study results are reviewed, and the major error sources described. Tracking and data fitting strategy actually used in the real time operations is itemized, and Deep Space Network data available for orbit fitting during the mission and the auxiliary information used by the navigation team are described. A detailed orbit fitting history of the first four revolutions of the satellite orbit of Mariner 9 is presented, with emphasis on the convergence problems and the delivered solution for the first orbit trim maneuver. Also included are a solution accuracy summary, the history of the spacecraft orbit osculating elements, the results of verifying the radio solutions with TV imaging data, and a summary of the normal points generated for the relativity experiment.
Genetic Algorithm for Initial Orbit Determination with Too Short Arc (Continued)
NASA Astrophysics Data System (ADS)
Li, X. R.; Wang, X.
2016-03-01
When using the genetic algorithm to solve the problem of too-short-arc (TSA) determination, due to the difference of computing processes between the genetic algorithm and classical method, the methods for outliers editing are no longer applicable. In the genetic algorithm, the robust estimation is acquired by means of using different loss functions in the fitness function, then the outlier problem of TSAs is solved. Compared with the classical method, the application of loss functions in the genetic algorithm is greatly simplified. Through the comparison of results of different loss functions, it is clear that the methods of least median square and least trimmed square can greatly improve the robustness of TSAs, and have a high breakdown point.
Lunar Reconnaissance Orbiter Orbit Determination Accuracy Analysis
NASA Technical Reports Server (NTRS)
Slojkowski, Steven E.
2014-01-01
Results from operational OD produced by the NASA Goddard Flight Dynamics Facility for the LRO nominal and extended mission are presented. During the LRO nominal mission, when LRO flew in a low circular orbit, orbit determination requirements were met nearly 100% of the time. When the extended mission began, LRO returned to a more elliptical frozen orbit where gravity and other modeling errors caused numerous violations of mission accuracy requirements. Prediction accuracy is particularly challenged during periods when LRO is in full-Sun. A series of improvements to LRO orbit determination are presented, including implementation of new lunar gravity models, improved spacecraft solar radiation pressure modeling using a dynamic multi-plate area model, a shorter orbit determination arc length, and a constrained plane method for estimation. The analysis presented in this paper shows that updated lunar gravity models improved accuracy in the frozen orbit, and a multiplate dynamic area model improves prediction accuracy during full-Sun orbit periods. Implementation of a 36-hour tracking data arc and plane constraints during edge-on orbit geometry also provide benefits. A comparison of the operational solutions to precision orbit determination solutions shows agreement on a 100- to 250-meter level in definitive accuracy.
NASA Technical Reports Server (NTRS)
Carpenter, James R.; Berry, Kevin; Gregpru. Late; Speckman, Keith; Hur-Diaz, Sun; Surka, Derek; Gaylor, Dave
2010-01-01
The Orbit Determination Toolbox is an orbit determination (OD) analysis tool based on MATLAB and Java that provides a flexible way to do early mission analysis. The toolbox is primarily intended for advanced mission analysis such as might be performed in concept exploration, proposal, early design phase, or rapid design center environments. The emphasis is on flexibility, but it has enough fidelity to produce credible results. Insight into all flight dynamics source code is provided. MATLAB is the primary user interface and is used for piecing together measurement and dynamic models. The Java Astrodynamics Toolbox is used as an engine for things that might be slow or inefficient in MATLAB, such as high-fidelity trajectory propagation, lunar and planetary ephemeris look-ups, precession, nutation, polar motion calculations, ephemeris file parsing, and the like. The primary analysis functions are sequential filter/smoother and batch least-squares commands that incorporate Monte-Carlo data simulation, linear covariance analysis, measurement processing, and plotting capabilities at the generic level. These functions have a user interface that is based on that of the MATLAB ODE suite. To perform a specific analysis, users write MATLAB functions that implement truth and design system models. The user provides his or her models as inputs to the filter commands. The software provides a capability to publish and subscribe to a software bus that is compliant with the NASA Goddard Mission Services Evolution Center (GMSEC) standards, to exchange data with other flight dynamics tools to simplify the flight dynamics design cycle. Using the publish and subscribe approach allows for analysts in a rapid design center environment to seamlessly incorporate changes in spacecraft and mission design into navigation analysis and vice versa.
Shadowing Lemma and chaotic orbit determination
NASA Astrophysics Data System (ADS)
Spoto, Federica; Milani, Andrea
2016-03-01
Orbit determination is possible for a chaotic orbit of a dynamical system, given a finite set of observations, provided the initial conditions are at the central time. The Shadowing Lemma (Anosov 1967; Bowen in J Differ Equ 18:333-356, 1975) can be seen as a way to connect the orbit obtained using the observations with a real trajectory. An orbit is a shadowing of the trajectory if it stays close to the real trajectory for some amount of time. In a simple discrete model, the standard map, we tackle the problem of chaotic orbit determination when observations extend beyond the predictability horizon. If the orbit is hyperbolic, a shadowing orbit is computed by the least squares orbit determination. We test both the convergence of the orbit determination iterative procedure and the behaviour of the uncertainties as a function of the maximum number of map iterations observed. When the initial conditions belong to a chaotic orbit, the orbit determination is made impossible by numerical instability beyond a computability horizon, which can be approximately predicted by a simple formula. Moreover, the uncertainty of the results is sharply increased if a dynamical parameter is added to the initial conditions as parameter to be estimated. The Shadowing Lemma does not dictate what the asymptotic behaviour of the uncertainties should be. These phenomena have significant implications, which remain to be studied, in practical problems of orbit determination involving chaos, such as the chaotic rotation state of a celestial body and a chaotic orbit of a planet-crossing asteroid undergoing many close approaches.
Kaguya Orbit Determination from JPL
NASA Technical Reports Server (NTRS)
Haw, Robert J.; Mottinger, N. A.; Graat, E. J.; Jefferson, D. C.; Park, R.; Menom, P.; Higa, E.
2008-01-01
Selene (re-named 'Kaguya' after launch) is an unmanned mission to the Moon navigated, in part, by JPL personnel. Launched by an H-IIA rocket on September 14, 2007 from Tanegashima Space Center, Kaguya entered a high, Earth-centered phasing orbit with apogee near the radius of the Moon's orbit. After 19 days and two orbits of Earth, Kaguya entered lunar orbit. Over the next 2 weeks the spacecraft decreased its apolune altitude until reaching a circular, 100 kilometer altitude orbit. This paper describes NASA/JPL's participation in the JAXA/Kaguya mission during that 5 week period, wherein JPL provided tracking data and orbit determination support for Kaguya.
Orbit Determination of the Lunar Reconnaissance Orbiter
NASA Technical Reports Server (NTRS)
Mazarico, Erwan; Rowlands, D. D.; Neumann, G. A.; Smith, D. E.; Torrence, M. H.; Lemoine, F. G.; Zuber, M. T.
2011-01-01
We present the results on precision orbit determination from the radio science investigation of the Lunar Reconnaissance Orbiter (LRO) spacecraft. We describe the data, modeling and methods used to achieve position knowledge several times better than the required 50-100m (in total position), over the period from 13 July 2009 to 31 January 2011. In addition to the near-continuous radiometric tracking data, we include altimetric data from the Lunar Orbiter Laser Altimeter (LOLA) in the form of crossover measurements, and show that they strongly improve the accuracy of the orbit reconstruction (total position overlap differences decrease from approx.70m to approx.23 m). To refine the spacecraft trajectory further, we develop a lunar gravity field by combining the newly acquired LRO data with the historical data. The reprocessing of the spacecraft trajectory with that model shows significantly increased accuracy (approx.20m with only the radiometric data, and approx.14m with the addition of the altimetric crossovers). LOLA topographic maps and calibration data from the Lunar Reconnaissance Orbiter Camera were used to supplement the results of the overlap analysis and demonstrate the trajectory accuracy.
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.
Precise Orbit Determination for ALOS
NASA Technical Reports Server (NTRS)
Nakamura, Ryo; Nakamura, Shinichi; Kudo, Nobuo; Katagiri, Seiji
2007-01-01
The Advanced Land Observing Satellite (ALOS) has been developed to contribute to the fields of mapping, precise regional land coverage observation, disaster monitoring, and resource surveying. Because the mounted sensors need high geometrical accuracy, precise orbit determination for ALOS is essential for satisfying the mission objectives. So ALOS mounts a GPS receiver and a Laser Reflector (LR) for Satellite Laser Ranging (SLR). This paper deals with the precise orbit determination experiments for ALOS using Global and High Accuracy Trajectory determination System (GUTS) and the evaluation of the orbit determination accuracy by SLR data. The results show that, even though the GPS receiver loses lock of GPS signals more frequently than expected, GPS-based orbit is consistent with SLR-based orbit. And considering the 1 sigma error, orbit determination accuracy of a few decimeters (peak-to-peak) was achieved.
Spitzer Orbit Determination During In-orbit Checkout Phase
NASA Technical Reports Server (NTRS)
Menon, Premkumar R.
2004-01-01
The Spitzer Space Telescope was injected into heliocentric orbit on August 25, 2003 to observe and study astrophysical phenomena in the infrared range of frequencies. The initial 60 days was dedicated to Spitzer's "In-Orbit Checkout (IOC)" efforts. During this time high levels of Helium venting were used to cool down the telescope. Attitude control was done using reaction wheels, which in turn were de-saturated using cold gas Nitrogen thrusting. Dense tracking data (nearly continuous) by the Deep Space network (DSN) were used to perform orbit determination and to assess any possible venting imbalance. Only Doppler data were available for navigation. This paper deals with navigation efforts during the IOC phase. It includes Dust Cover Ejection (DCE) monitoring, orbit determination strategy validation and results and assessment of non-gravitational accelerations acting on Spitzer including that due to possible imbalance in Helium venting.
Orbit Determination Issues for Libration Point Orbits
NASA Technical Reports Server (NTRS)
Beckman, Mark; Bauer, Frank (Technical Monitor)
2002-01-01
Libration point mission designers require knowledge of orbital accuracy for a variety of analyses including station keeping control strategies, transfer trajectory design, and formation and constellation control. Past publications have detailed orbit determination (OD) results from individual libration point missions. This paper collects both published and unpublished results from four previous libration point missions (ISEE (International Sun-Earth Explorer) -3, SOHO (Solar and Heliospheric Observatory), ACE (Advanced Composition Explorer) and MAP (Microwave Anisotropy Probe)) supported by Goddard Space Flight Center's Guidance, Navigation & Control Center. The results of those missions are presented along with OD issues specific to each mission. All past missions have been limited to ground based tracking through NASA ground sites using standard range and Doppler measurement types. Advanced technology is enabling other OD options including onboard navigation using seaboard attitude sensors and the use of the Very Long Baseline Interferometry (VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options potentially enable missions to reduce coherent dedicated tracking passes while maintaining orbital accuracy. With the increased projected loading of the DSN (Deep Space Network), missions must find alternatives to the standard OD scenario.
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.
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.
Geostationary orbit determination using SATRE
NASA Astrophysics Data System (ADS)
Lei, Hui; Li, ZhiGang; Yang, XuHai; Wu, WenJun; Cheng, Xuan; Yang, Ying; Feng, ChuGang
2011-09-01
A new strategy of precise orbit determination (POD) for GEO (Geostationary Earth Orbit) satellite using SATRE (SAtellite Time and Ranging Equipment) is presented. Two observation modes are proposed and different channels of the same instruments are used to construct different observation modes, one mode receiving time signals from their own station and the other mode receiving time signals from each other for two stations called pairs of combined observations. Using data from such a tracking network in China, the results for both modes are compared. The precise orbit determination for the Sino-1 satellite using the data from 6 June 2005 to 13 June 2005 has been carried out in this work. The RMS (Root-Mean-Square) of observing residuals for 3-day solutions with the former mode is better than 9.1 cm. The RMS of observing residuals for 3-day solutions with the latter mode is better than 4.8 cm, much better than the former mode. Orbital overlapping (3-day orbit solution with 1-day orbit overlap) tests show that the RMS of the orbit difference for the former mode is 0.16 m in the radial direction, 0.53 m in the along-track direction, 0.97 m in the cross-track direction and 1.12 m in the 3-dimension position and the RMS of the orbit difference for the latter mode is 0.36 m in the radial direction, 0.89 m in the along-track direction, 1.18 m in the cross-track direction and 1.52 m in the 3-dimension position, almost the same as the former mode. All the experiments indicate that a meter-level accuracy of orbit determination for geostationary satellite is achievable.
The determination of orbits using Picard iteration
NASA Technical Reports Server (NTRS)
Mikkilineni, R. P.; Feagin, T.
1975-01-01
The determination of orbits by using Picard iteration is reported. This is a direct extension of the classical method of Picard that has been used in finding approximate solutions of nonlinear differential equations for a variety of problems. The application of the Picard method of successive approximations to the initial value and the two point boundary value problems is given.
Lunar Reconnaissance Orbiter Orbit Determination Accuracy Analysis
NASA Technical Reports Server (NTRS)
Slojkowski, Steven E.
2014-01-01
LRO definitive and predictive accuracy requirements were easily met in the nominal mission orbit, using the LP150Q lunar gravity model. center dot Accuracy of the LP150Q model is poorer in the extended mission elliptical orbit. center dot Later lunar gravity models, in particular GSFC-GRAIL-270, improve OD accuracy in the extended mission. center dot Implementation of a constrained plane when the orbit is within 45 degrees of the Earth-Moon line improves cross-track accuracy. center dot Prediction accuracy is still challenged during full-Sun periods due to coarse spacecraft area modeling - Implementation of a multi-plate area model with definitive attitude input can eliminate prediction violations. - The FDF is evaluating using analytic and predicted attitude modeling to improve full-Sun prediction accuracy. center dot Comparison of FDF ephemeris file to high-precision ephemeris files provides gross confirmation that overlap compares properly assess orbit accuracy.
Gravity Probe B orbit determination
NASA Astrophysics Data System (ADS)
Shestople, P.; Ndili, A.; Hanuschak, G.; Parkinson, B. W.; Small, H.
2015-11-01
The Gravity Probe B (GP-B) satellite was equipped with a pair of redundant Global Positioning System (GPS) receivers used to provide navigation solutions for real-time and post-processed orbit determination (OD), as well as to establish the relation between vehicle time and coordinated universal time. The receivers performed better than the real-time position requirement of 100 m rms per axis. Post-processed solutions indicated an rms position error of 2.5 m and an rms velocity error of 2.2 mm s-1. Satellite laser ranging measurements provided independent verification of the GPS-derived GP-B orbit. We discuss the modifications and performance of the Trimble Advance Navigation System Vector III GPS receivers. We describe the GP-B precision orbit and detail the OD methodology, including ephemeris errors and the laser ranging measurements.
Single Frequency GPS Orbit Determination for Low Earth Orbiters
NASA Technical Reports Server (NTRS)
Bertiger, Willy; Wu, Sien-Chong
1996-01-01
A number of missions in the future are planning to use GPS for precision orbit determination. Cost considerations and receiver availability make single frequency GPS receivers attractive if the orbit accuracy requirements can be met.
Prestellar cores: initial orbit and boundary
NASA Astrophysics Data System (ADS)
Horedt, G. P.
2016-06-01
The initial orbit of a prestellar core in the resisting intercore medium is found to be an elliptic spiral round the mass centre of the parent molecular cloud (clump), with exponentially decreasing semiaxes, high constant eccentricity, and constant period. Prestellar cores are stable against perturbations caused by the parent cloud, if the corresponding mean density contrast is larger than about 10. This value defines the safe boundary of a prestellar core within its parent cloud and is in accordance with observations.
Low thrust orbit determination program
NASA Technical Reports Server (NTRS)
Hong, P. E.; Shults, G. L.; Huling, K. R.; Ratliff, C. W.
1972-01-01
Logical flow and guidelines are provided for the construction of a low thrust orbit determination computer program. The program, tentatively called FRACAS (filter response analysis for continuously accelerating spacecraft), is capable of generating a reference low thrust trajectory, performing a linear covariance analysis of guidance and navigation processes, and analyzing trajectory nonlinearities in Monte Carlo fashion. The choice of trajectory, guidance and navigation models has been made after extensive literature surveys and investigation of previous software. A key part of program design relied upon experience gained in developing and using Martin Marietta Aerospace programs: TOPSEP (Targeting/Optimization for Solar Electric Propulsion), GODSEP (Guidance and Orbit Determination for SEP) and SIMSEP (Simulation of SEP).
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.
The Seasat Precision Orbit Determination Experiment
NASA Technical Reports Server (NTRS)
Tapley, B. D.; Born, G. H.
1980-01-01
The objectives and conclusions reached during the Seasat Precision Orbit Determination Experiment are discussed. It is noted that the activities of the experiment team included extensive software calibration and validation and an intense effort to validate and improve the dynamic models which describe the satellite's motion. Significant improvement in the gravitational model was obtained during the experiment, and it is pointed out that the current accuracy of the Seasat altitude ephemeris is 1.5 m rms. An altitude ephemeris for the Seasat spacecraft with an accuracy of 0.5 m rms is seen as possible with further improvements in the geopotential, atmospheric drag, and solar radiation pressure models. It is concluded that since altimetry missions with a 2-cm precision altimeter are contemplated, the precision orbit determination effort initiated under the Seasat Project must be continued and expanded.
Information Measures for Statistical Orbit Determination
ERIC Educational Resources Information Center
Mashiku, Alinda K.
2013-01-01
The current Situational Space Awareness (SSA) is faced with a huge task of tracking the increasing number of space objects. The tracking of space objects requires frequent and accurate monitoring for orbit maintenance and collision avoidance using methods for statistical orbit determination. Statistical orbit determination enables us to obtain…
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
NASA Astrophysics Data System (ADS)
Shakun, L. S.; Koshkin, N. I.
2014-06-01
The number of artificial space objects in the low Earth orbit has been continuously increasing. That raises the requirements for the accuracy of measurement of their coordinates and for the precision of the prediction of their motion. The accuracy of the prediction can be improved if the actual current orientation of the non-spherical satellite is taken into account. In so doing, it becomes possible to directly determine the atmospheric density along the orbit. The problem solution is to regularly conduct the photometric surveillances of a large number of satellites and monitor the parameters of their rotation around the centre of mass. To do that, it is necessary to get and promptly process large video arrays, containing pictures of a satellite against the background stars. In the present paper, the method for the simultaneous measurement of coordinates and brightness of the low Earth orbit space objects against the background stars when they are tracked by telescope KT-50 with the mirror diameter of 50 cm and with video camera WAT-209H2 is considered. The problem of determination of the moments of exposures of images is examined in detail. The estimation of the accuracy of measuring both the apparent coordinates of stars and their photometry is given on the example of observation of the open star cluster. In the presented observations, the standard deviation of one position measured is 1σ, the accuracy of determination of the moment of exposure of images is better than 0.0001 s. The estimate of the standard deviation of one measurement of brightness is 0.1m. Some examples of the results of surveillances of satellites are also presented in the paper.
Orbit determination by genetic algorithm and application to GEO observation
NASA Astrophysics Data System (ADS)
Hinagawa, Hideaki; Yamaoka, Hitoshi; Hanada, Toshiya
2014-02-01
This paper demonstrates an initial orbit determination method that solves the problem by a genetic algorithm using two well-known solutions for the Lambert's problem: universal variable method and Battin method. This paper also suggests an intuitive error evaluation method in terms of rotational angle and orbit shape by separating orbit elements into two groups. As reference orbit, mean orbit elements (original two-lines elements) and osculating orbit elements considering the J2 effect are adopted and compared. Our proposed orbit determination method has been tested with actual optical observations of a geosynchronous spacecraft. It should be noted that this demonstration of the orbit determination is limited to one test case. This observation was conducted during approximately 70 min on 2013/05/15 UT. Our method was compared with the orbit elements propagated by SGP4 using the TLE of the spacecraft. The result indicates that our proposed method had a slightly better performance on estimating orbit shape than Gauss's methods and Escobal's method by 120 km. In addition, the result of the rotational angle is closer to the osculating orbit elements than the mean orbit elements by 0.02°, which supports that the estimated orbit is valid.
Orbit Determination System for Low Earth Orbit Satellites
NASA Technical Reports Server (NTRS)
Elisha, Yossi; Shyldkrot, Haim; Hankin, Maxim
2007-01-01
The IAI/MBT Precise Orbit Determination system for Low Earth Orbit satellites is presented. The system is based on GPS pesudorange and carrier phase measurements and implements the Reduced Dynamics method. The GPS measurements model, the dynamic model, and the least squares orbit determination are discussed. Results are shown for data from the CHAMP satellite and for simulated data from the ROKAR GPS receiver. In both cases the one sigma 3D position and velocity accuracy is about 0.2 m and 0.5 mm/sec respectively.
Precision Orbit Determination for the Lunar Reconnaissance Orbiter
NASA Astrophysics Data System (ADS)
Lemoine, F. G.; Mazarico, E.; Rowlands, D. D.; Torrence, M. H.; McGarry, J. F.; Neumann, G. A.; Mao, D.; Smith, D. E.; Zuber, M. T.
2010-05-01
The Lunar Reconnaissance Orbiter (LRO) spacecraft was launched on June 18, 2009. In mid-September 2009, the spacecraft orbit was changed from its commissioning orbit (30 x 216 km polar) to a quasi-frozen polar orbit with an average altitude of 50km (+-15km). One of the goals of the LRO mission is to develop a new lunar reference frame to facilitate future exploration. Precision Orbit Determination is used to achieve the accuracy requirements, and to precisely geolocate the high-resolution datasets obtained by the LRO instruments. In addition to the tracking data most commonly used to determine spacecraft orbits in planetary missions (radiometric Range and Doppler), LRO benefits from two other types of orbital constraints, both enabled by the Lunar Orbiter Laser Altimeter (LOLA) instrument. The altimetric data collected as the instrument's primary purpose can be used to derive constraints on the orbit geometry at the times of laser groundtrack intersections (crossovers). The multi-beam configuration and high firing-rate of LOLA further improves the strength of these crossovers, compared to what was possible with the MOLA instrument onboard Mars Global Surveyor (MGS). Furthermore, one-way laser ranges (LR) between Earth International Laser Ranging Service (ILRS) stations and the spacecraft are made possible by the addition of a small telescope mounted on the spacecraft high-gain antenna. The photons received from Earth are transmitted to one LOLA detector by a fiber optics bundle. Thanks to the accuracy of the LOLA timing system, the precision of 5-s LR normal points is below 10cm. We present the first results of the Precision Orbit Determination (POD) of LRO through the commissioning and nominal phases of the mission. Orbit quality is discussed, and various gravity fields are evaluated with the new (independent) LRO radio tracking data. The altimetric crossovers are used as an independent data type to evaluate the quality of the orbits. The contribution of the LR
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.
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 Δv 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.
Determination of GPS orbits to submeter accuracy
NASA Technical Reports Server (NTRS)
Bertiger, W. I.; Lichten, S. M.; Katsigris, E. C.
1988-01-01
Orbits for satellites of the Global Positioning System (GPS) were determined with submeter accuracy. Tests used to assess orbital accuracy include orbit comparisons from independent data sets, orbit prediction, ground baseline determination, and formal errors. One satellite tracked 8 hours each day shows rms error below 1 m even when predicted more than 3 days outside of a 1-week data arc. Differential tracking of the GPS satellites in high Earth orbit provides a powerful relative positioning capability, even when a relatively small continental U.S. fiducial tracking network is used with less than one-third of the full GPS constellation. To demonstrate this capability, baselines of up to 2000 km in North America were also determined with the GPS orbits. The 2000 km baselines show rms daily repeatability of 0.3 to 2 parts in 10 to the 8th power and agree with very long base interferometry (VLBI) solutions at the level of 1.5 parts in 10 to the 8th power. This GPS demonstration provides an opportunity to test different techniques for high-accuracy orbit determination for high Earth orbiters. The best GPS orbit strategies included data arcs of at least 1 week, process noise models for tropospheric fluctuations, estimation of GPS solar pressure coefficients, and combine processing of GPS carrier phase and pseudorange data. For data arc of 2 weeks, constrained process noise models for GPS dynamic parameters significantly improved the situation.
Precise Orbit Determination for Altimeter Satellites
NASA Astrophysics Data System (ADS)
Zelensky, N. P.; Luthcke, S. B.; Rowlands, D. D.; Lemoine, F. G.; Beckley, B. B.; Wang, Y.; Chinn, D. S.
2002-05-01
Orbit error remains a critical component in the error budget for all radar altimeter missions. This paper describes the ongoing work at GSFC to improve orbits for three radar altimeter satellites: TOPEX/POSEIDON (T/P), Jason, and Geosat Follow-On (GFO). 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 (2-3 cm radially) produced at GSFC. Jason, the T/P follow-on, is intended to continue measurement of the ocean surface with the same, if not better accuracy. Reaching the Jason centimeter accuracy orbit goal would greatly benefit the knowledge of ocean circulation. Several new POD strategies which promise significant improvement to the current T/P orbit are evaluated over one year of data. Also, preliminary, but very promising Jason POD results are presented. Orbit improvement for GFO has been dramatic, and has allowed this mission to provide a POESEIDON class altimeter product. The GFO Precise Orbit Ephemeris (POE) orbits are based on satellite laser ranging (SLR) tracking supplemented with GFO/GFO altimeter crossover data. The accuracy of these orbits were evaluated using several tests, including independent TOPEX/GFO altimeter crossover data. The orbit improvements are shown over the years 2000 and 2001 for which the POEs have been completed.
NASA Astrophysics Data System (ADS)
Xu, Peiliang
2015-04-01
Satellite orbits have been routinely used to produce models of the Earth's gravity field. The numerical integration method is most widely used by almost all major institutions to determine standard gravity models from space geodetic measurements. As a basic component of the method, the partial derivatives of a satellite orbit with respect to the force parameters to be determined, namely, the unknown harmonic coefficients of the gravitational model, have been first computed by setting the initial values of partial derivatives to zero. In this talk, we first design some simple mathematical examples to show that setting the initial values of partial derivatives to zero is generally erroneous mathematically. We then prove that it is prohibited physically. In other words, setting the initial values of partial derivatives to zero violates the physics of motion of celestial bodies. To conclude, the numerical integration method, as is widely used today by major institutions to produce standard satellite gravity models, is simply incorrect mathematically. As a direct consequence, further work is required to confirm whether the numerical integration method can still be used as a mathematical foundation to produce standard satellite gravity models. More details can be found in Xu (2009, Sci China Ser D-Earth Sci, 52, 562-566).
Lunar orbiter ranging data: initial results.
Mulholland, J D; Sjogren, W L
1967-01-01
Data from two Lunar Orbiter spacecraft have been used to test the significance of corrections to the lunar ephemeris. Range residuals of up to 1700 meters were reduced by an order of magnitude by application of the corrections, with most of the residuals reduced to less than 100 meters. Removal of gross errors in the ephemeris reveals residual patterns that may indicate errors in location of observing stations, as well as the expected effects of Lunar nonsphericity. PMID:17799149
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
Satellite orbit determination from an airborne platform
NASA Astrophysics Data System (ADS)
Shepard, M. M.; Foshee, J. J.
This paper describes the requirements, approach, and problems associated with autonomous satellite orbit determination from an airborne platform. The ability to perform orbit determination from an airborne platform removes the reliance on ground control facilities. Aircraft orbit determination offers a more robust system in that it is less susceptible to direct attack, sabotage, or nuclear disaster. Ranging on a satellite and the processing of range/range-rate data along with INS inputs to produce a set of orbital parameters to be transmitted to user terminals are discussed. Several algorithms that could be utilized by the user terminal to recover the satellite position/velocity data from the transmitted message are presented. The ability to compress the ephemeris message to a small size while remaining autonomous for a long period of time, as would be needed in future military communication satellites, is discussed.
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.
Two-tangent-impulse flyby of space target from an elliptic initial orbit
NASA Astrophysics Data System (ADS)
Zhu, Zhenglong; Yan, Ye
2016-05-01
The problem of two-tangent-impulse space target flyby is drawn in this paper. Unlike the traditional method, the transfer orbit is divided into two parts, i.e. the phasing orbit and the approaching orbit, considering the temporal and the spatial constraint respectively. The method of the phasing orbit and approaching orbit determination is formulated firstly. Then, the attainable target areas from a given position on the initial elliptic orbit, the feasible impulse position for approaching to a given point in space and the optimal approaching orbit to a given point are analyzed. Based on the optimal approaching orbit, the optimal solution for target flyby is provided in the time free case, and a suboptimal solution for the case of time-constrained. Numerical simulations demonstrate the application of the solutions to space target flyby in different scenarios.
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.
The GEOS-3 orbit determination investigation
NASA Technical Reports Server (NTRS)
Pisacane, V. L.; Eisner, A.; Yionoulis, S. M.; Mcconahy, R. J.; Black, H. D.; Pryor, L. L.
1978-01-01
The nature and improvement in satellite orbit determination when precise altimetric height data are used in combination with conventional tracking data was determined. A digital orbit determination program was developed that could singly or jointly use laser ranging, C-band ranging, Doppler range difference, and altimetric height data. Two intervals were selected and used in a preliminary evaluation of the altimeter data. With the data available, it was possible to determine the semimajor axis and eccentricity to within several kilometers, in addition to determining an altimeter height bias. When used jointly with a limited amount of either C-band or laser range data, it was shown that altimeter data can improve the orbit solution.
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 Ω, 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
NASA Technical Reports Server (NTRS)
Yee, C. P.; Kelbel, D. A.; Lee, T.; Dunham, J. B.; Mistretta, G. D.
1990-01-01
The influence of ionospheric refraction on orbit determination was studied through the use of the Orbit Determination Error Analysis System (ODEAS). The results of a study of the orbital state estimate errors due to the ionospheric refraction corrections, particularly for measurements involving spacecraft-to-spacecraft tracking links, are presented. In current operational practice at the Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF), the ionospheric refraction effects on the tracking measurements are modeled in the Goddard Trajectory Determination System (GTDS) using the Bent ionospheric model. While GTDS has the capability of incorporating the ionospheric refraction effects for measurements involving ground-to-spacecraft tracking links, such as those generated by the Ground Spaceflight Tracking and Data Network (GSTDN), it does not have the capability to incorporate the refraction effects for spacecraft-to-spacecraft tracking links for measurements generated by the Tracking and Data Relay Satellite System (TDRSS). The lack of this particular capability in GTDS raised some concern about the achievable accuracy of the estimated orbit for certain classes of spacecraft missions that require high-precision orbits. Using an enhanced research version of GTDS, some efforts have already been made to assess the importance of the spacecraft-to-spacecraft ionospheric refraction corrections in an orbit determination process. While these studies were performed using simulated data or real tracking data in definitive orbit determination modes, the study results presented here were obtained by means of covariance analysis simulating the weighted least-squares method used in orbit determination.
Filtering theory applied to orbit determination
NASA Technical Reports Server (NTRS)
Torroglosa, V.
1973-01-01
Modifications to the extended Kalman filter and the Jazwinski filter are made and compared with the classical extended Kalman filter in applications to orbit determination using real data. The results show that with the kind of data available today, the application of filtering theories in this field presents many problems.
Preliminary orbit determination for lunar satellites.
NASA Technical Reports Server (NTRS)
Lancaster, E. R.
1973-01-01
Methods for the determination of orbits of artificial lunar satellites from earth-based range rate measurements developed by Koskela (1964) and Bateman et al. (1966) are simplified and extended to include range measurements along with range rate measurements. For illustration, a numerical example is presented.
Algorithms for Autonomous GS Orbit Determination and Formation Flying
NASA Technical Reports Server (NTRS)
Moreau, Michael C.; Speed, Eden Denton-Trost; Axelrad, Penina; Leitner, Jesse (Technical Monitor)
2001-01-01
This final report for our study of autonomous Global Positioning System (GPS) satellite orbit determination comprises two sections. The first is the Ph.D. dissertation written by Michael C. Moreau entitled, "GPS Receiver Architecture for Autonomous Navigation in High Earth Orbits." Dr. Moreau's work was conducted under both this project and a NASA GSRP. His dissertation describes the key design features of a receiver specifically designed for autonomous operation in high earth orbits (HEO). He focused on the implementation and testing of these features for the GSFC PiVoT receiver. The second part is a memo describing a robust method for autonomous initialization of the orbit estimate given very little a priori information and sparse measurements. This is a key piece missing in the design of receivers for HEO.
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.
GPS-LEO orbiter occultation orbital analyses and event determination
NASA Astrophysics Data System (ADS)
Abdul Rashid, Z. A.; Cheng, P. P.
2003-04-01
A good knowledge of the vertical profiles of temperature and humidity throughout the atmosphere are crucial to understand the present state of the Earth's atmosphere and it's modeling. The application of radio occultation technique has a heritage of over 2 decades in NASA's planetary exploration program to study the atmosphere of most of the major bodies in the solar system. Results from NASA's planetary program experiment have proven to be very effective at characterizing the atmosphere of a planet. However, the use of radio occultation technique to remote sensing the Earth's atmosphere is only practical to be implemented recently with the advent of the matured Global Positioning System (GPS). The GPS occultation technique is well suited to observe the Earth's atmosphere, due to it excellent geographical coverage, all weather capability, long-term stability, self-calibration and high vertical resolution. The GPS/MET (GPS Meteorology) experiment launched in April 1995 is the proof-of-concept of this technique. The results from this experiment is appealing and shown that the GPS occultation technique is a promising candidate to monitor the Earth's atmosphere. With the advancement of receiver technologies and lower system cost, the GPS occultation technique is a promising tool to predict the long-term climatic changes and numerical weather modeling of the Earth's atmosphere at a higher precision. This paper briefly describes the radio occultation concept and the GPS satellite systems, which form the basis understanding of this subject matter. This is followed by a detail description of the occultation geometries between the GPS satellites and a LEO orbiter. A method to determine the occultation event is discussed and thoroughly analyzed in terms of orbit inclinations, altitudes, receiver sampling rates, antenna positioning (aft and fore pointing), and antenna mask angles. A simulator is developed using MATLAB for the orbital analyses and occultation determination in
Accelerometers for Precise GNSS Orbit Determination
NASA Astrophysics Data System (ADS)
Hugentobler, Urs; Schlicht, Anja
2016-07-01
The solar radiation pressure is the largest non-gravitational acceleration on GNSS satellites limiting the accuracy of precise orbit models. Other non-gravitational accelerations may be thrusts for station keeping maneuvers. Accelerometers measure the motion of a test mass that is shielded against satellite surface forces with respect to a cage that is rigidly connected to the satellite. They can thus be used to measure these difficult-to-model non-gravitational accelerations. Accelerometers however typically show correlated noise as well as a drift of the scaling factors converting measured voltages to accelerations. The scaling thus needs to be regularly calibrated. The presented study is based on several simulated scenarios including orbit determination of accelerometer-equipped Galileo satellites. It shall evaluate different options on how to accommodate accelerometer measurements in the orbit integrator, indicate to what extent currently available accelerometers can be used to improve the modeling of non-gravitational accelerations on GNSS satellites for precise orbit determination, and assess the necessary requirements for an accelerometer that can serve this purpose.
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.
Formation Flying In Highly Elliptical Orbits Initializing the Formation
NASA Technical Reports Server (NTRS)
Mailhe, Laurie; Schiff, Conrad; Hughes, Steven
2000-01-01
In this paper several methods are examined for initializing formations in which all spacecraft start in a common elliptical orbit subsequent to separation from the launch vehicle. The tetrahedron formation used on missions such as the Magnetospheric Multiscale (MMS), Auroral Multiscale Midex (AMM), and Cluster is used as a test bed Such a formation provides full three degrees-of-freedom in the relative motion about the reference orbit and is germane to several missions. The type of maneuver strategy that can be employed depends on the specific initial conditions of each member of the formation. Single-impulse maneuvers based on a Gaussian variation-of-parameters (VOP) approach, while operationally simple and intuitively-based, work only in a limited sense for a special class of initial conditions. These 'tailored' initial conditions are characterized as having only a few of the Keplerian elements different from the reference orbit. Attempts to achieve more generic initial conditions exceed the capabilities of the single impulse VOP. For these cases, multiple-impulse implementations are always possible but are generally less intuitive than the single-impulse case. The four-impulse VOP formalism discussed by Schaub is examined but smaller delta-V costs are achieved in our test problem by optimizing a Lambert solution.
Using Onboard Telemetry for MAVEN Orbit Determination
NASA Technical Reports Server (NTRS)
Lam, Try; Trawny, Nikolas; Lee, Clifford
2013-01-01
Determination of the spacecraft state has been traditional done using radiometric tracking data before and after the atmosphere drag pass. This paper describes our approach and results to include onboard telemetry measurements in addition to radiometric observables to refine the reconstructed trajectory estimate for the Mars Atmosphere and Volatile Evolution Mission (MAVEN). Uncertainties in the Mars atmosphere models, combined with non-continuous tracking degrade navigation accuracy, making MAVEN a key candidate for using onboard telemetry data to help complement its orbit determination process.
Parallel Computation of Orbit Determination for Space Debris Population
NASA Astrophysics Data System (ADS)
Olmedo, Estrella; Sanchez-Ortiz, Noelia; Ramos-Lerate, Mercedes
2009-03-01
In this work we present an algorithm for computing Orbit Determination for Space Debris population. The method presents a high degree of parallelism. That means that the number of available computers divides the computational effort. The context of this work and the later scope is to have the capability of cataloguing and correlating the Space Debris population. In this sense, as better the accuracy provided by the orbit determination is, more accurate will be the estimation of the state vectors corresponding to the debris objects and better will be the accuracy of the future catalogue of Space Debris. As more objects we can determinate the corresponding orbit, more complete will be the future catalogue. Therefore numerical tools for orbit determination are a key point in the development of a future ESSAS. The first time that a new object is observed, six measurements (these measurements may come from RADAR, Ground Based Telescope or Space Based Telescope) are required for computing an Initial Orbit Determination (IOD). After that, the Initial Estimated State Vector (IESV) is improved within the next-coming measurement. The idea of this method is the following. From six initial measurements, we compute the IOD following the same ideas of [1]. We compute also the initial knowledge covariance matrix (IKCM) corresponding to the IESV. In general, the numerical error of the IOD is too big for processing the following measurements with a conventional numerical filter (like the Square Root Information Filter (SRIF)). The problem is that the improvement of the accuracy in the IOD is not an easy task in those cases with large initial error. However the computed IKCM give a realistic approximation of the committed error in the IOD. The proposed algorithm uses the IKCM for generating a cloud of IESVs. All the IESV inside the cloud are processed with a new and much smaller IKCM by using SRIF. In such a way that the ones that are close enough to the real state vector (and thus
Orbit determination of Tance-1 satellite using VLBI data
NASA Astrophysics Data System (ADS)
Huang, Y.; Hu, X. G.; Huang, C.; Jiang, D. R.
2006-01-01
On 30 December, 2003, China successfully launched the first satellite Tance-1 of Chinese Geospace Double Star Exploration Program, i.e. "Double Star Program (DSP)", on an improved Long March 2C launch vehicle. The Tance-1 satellite is operating at an orbit around the earth with a 550km perigee, 78000km apogee and 28.5 degree inclination.VLBI technique can track Tance-1 satellite or even far satellites such as lunar vehicles. To validate the VLBI technique in the on-going Chinese lunar exploration mission, Shanghai Astronomical Observatory (SHAO) organized to track the Tance-1 satellite with Chinese three VLBI stations: Shanghai, Kunming and Urumchi Orbit Determination (OD) of the Tance-1 satellite with about two days VLBI dada, and the capability of OD with VLBI data are studied. The results show that the VLBI-based orbit solutions improve the fit level over the initial orbit. The VLBI-delay-based orbit solution shows that the RMS of residuals of VLBI delay data is about 5.5m, and about 2.0cm/s for the withheld VLBI delay rate data. The VLBI-delay-rate-based orbit solution shows that the RMS of residuals of VLBI delay rate data is about 1.3cm/s, and about 29m for the withheld VLBI delay data. In the situation of orbit determination with VLBI delay and delay rate data with data sigma 5.5m and 1.3cm/s respectively, the RMS of residuals are 5.5,m and 2.0cm/s respectively. The simulation data assess the performance of the solutions. Considering the dynamic model errors of the Tance-1 satellite, the accuracy of the position is about km magnitude, and the accuracy of the velocity is about cm/s magnitude. The simulation work also show the dramatic accuracy improvement of OD with VLBI and USB combined.
NASA Astrophysics Data System (ADS)
Mazarico, E.; Rowlands, D. D.; Neumann, G. A.; Lemoine, F. G.; Torrence, M. H.; Smith, D. E.; Zuber, M. T.; Mao, D.
2010-12-01
We present results of the Precision Orbit Determination work undertaken by the Lunar Orbiter Laser Altimeter (LOLA) Science Team for the Lunar Reconnaissance Orbiter (LRO) mission, in order to meet the position knowledge accuracy requirements (50-m total position) and to precisely geolocate the LRO datasets. In addition to the radiometric tracking data, one-way laser ranges (LR) between Earth stations and the spacecraft are made possible by a small telescope mounted on the spacecraft high-gain antenna. The photons received from Earth are transmitted to one LOLA detector by a fiber optics bundle. The LOLA timing system enables 5-s LR normal points with precision better than 10cm. Other types of geodetic constraints are derived from the altimetric data itself. The orbit geometry can be constrained at the times of laser groundtrack intersections (crossovers). Due to the Moon's slow rotation, orbit solutions and normal equations including altimeter crossovers are processed and created in one month batches. Recent high-resolution topographic maps near the lunar poles are used to produce a new kind of geodetic constraints. Purely geometric, those do not necessitate actual groundtrack intersections. We assess the contributions of those data types, and the quality of our orbits. Solutions which use altimetric crossover meet the horizontal 50-m requirement, and perform usually better (10-20m). We also obtain gravity field solutions based on LRO and historical data. The various LRO data are accumulated into normal equations, separately for each one month batch and for each measurement type, which enables the final weights to be adjusted during the least-squares inversion step. Expansion coefficients to degree and order 150 are estimated, and a Kaula rule is still needed to stabilize the farside field. The gravity field solutions are compared to previous solutions (GLGM-3, LP150Q, SGM100h) and the geopotential predicted from the latest LOLA spherical harmonic expansion.
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.
Mars exploration rovers orbit determination system modeling
NASA Astrophysics Data System (ADS)
Wawrzyniak, Geoffrey; Baird, Darren; Graat, Eric; McElrath, Tim; Portock, Brian; Watkins, Michael
2006-06-01
From June 2003 to January 2004, two spinning spacecraft journeyed from Earth to Mars. A team of navigators at the Jet Propulsion Laboratory (JPL) accurately determined the orbits of both Mars Exploration Rovers, Spirit and Opportunity. For the navigation process to be successful, the team needed to know how nongravitational effects and how measurement system properties affected the trajectory and data modeling. To accomplish this, in addition to the standard gravitational and radiometric modeling of the spacecraft, a calibration was performed on each spacecraft to determine the amount of ΔV that might occur during a turn, a high-fidelity solar-radiation-pressure model was created, the spin signature was removed from the tracking data, the station locations of the Deep Space Network were resurveyed, and a model of interplanetary charged particles was developed. The result of this effort was near-perfect accuracy, surpassing the tight atmospheric-entry requirements for navigation of both spacecraft.
OrbView-3 Initial On-Orbit Characterization
NASA Technical Reports Server (NTRS)
Ross, Kent; Blonski, Slawomir; Holekamp, Kara; Pagnutti, Mary; Zanoni, Vicki; Carver, David; Fendley, Debbie; Smith, Charles
2004-01-01
NASA at Stennis Space Center (SSC) established a Space Act Agreement with Orbital Sciences Corporation (OSC) and ORBIMAGE Inc. to collaborate on the characterization of the OrbView-3 system and its imagery products and to develop characterization techniques further. In accordance with the agreement, NASA performed an independent radiometric, spatial, and geopositional accuracy assessment of OrbView-3 imagery acquired before completion of the system's initial on-orbit checkout. OSC acquired OrbView-3 imagery over SSC from July 2003 through January 2004, and NASA collected ground reference information coincident with many of these acquisitions. After evaluating all acquisitions, NASA deemed two multispectral images and five panchromatic images useful for characterization. NASA then performed radiometric, spatial, and geopositional characterizations.
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.
NASA Astrophysics Data System (ADS)
Maier, A.; Baur, O.; Krauss, S.
2014-04-01
This contribution deals with Precise Orbit Determination of the Lunar Reconnaissance Orbiter, which is tracked with optical laser ranges in addition to radiometric Doppler range-rates and range observations. The optimum parameterization is assessed by overlap analysis tests that indicate the inner precision of the computed orbits. Information about the very long wavelengths of the lunar gravity field is inferred from the spacecraft positions. The NASA software packages GEODYN II and SOLVE were used for orbit determination and gravity field recovery [1].
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.
NASA Astrophysics Data System (ADS)
Murison, Marc A.
2006-06-01
This paper addresses the characterization of the precision of observationally determined orbit parameters when optical observations are taken of an artificial satellite ("target") from another orbiting body ("platform"). Of interest are, among others, optimal platform orbits and optimal observing strategies for a given level of observational astrometric precision and for certain types of target orbits. Classical orbit determination methods are not particularly amenable for gaining analytical insight into the characterization of the determined orbital parameter errors. Here we make an attempt to bypass classical orbit determination and look for an approach that can instead make use of certain approximations to the relative distance and velocity vectors. Furthermore, given the modern possibility for spectroscopic optical instruments in space, we also investigate what may additionally be gained from radial velocity observations. We start with the distance and velocity vectors of an orbiting target body with respect to an orbiting observation platform. We approximate the relative distance and velocity vectors, allowed by certain assumptions such as small eccentricities, relative inclination angle(s), and ratio of orbit radii. We then analytically propagate the observational errors through the equations and characterize what target orbit parameter errors we are able. It turns out this is more difficult than anticipated at first. We then perform numerical simulations to more completely characterize the behaviors of the determined orbit parameter errors.
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.
GPS as an orbit determination subsystems
NASA Technical Reports Server (NTRS)
Fennessey, Richard; Roberts, Pat; Knight, Robin; Vanvolkinburg, Bart
1995-01-01
This paper evaluates the use of Global Positioning System (GPS) receivers as a primary source of tracking data for low-Earth orbit satellites. GPS data is an alternative to using range, azimuth, elevation, and range-rate (RAER) data from the Air Force Satellite Control Network antennas, the Space Ground Link System (SGLS). This evaluation is applicable to missions such as Skipper, a joint U.S. and Russian atmosphere research mission, that will rely on a GPS receiver as a primary tracking data source. The Detachment 2, Space and Missile Systems Center's Test Support Complex (TSC) conducted the evaluation based on receiver data from the Space Test Experiment Platform Mission O (STEP-O) and Advanced Photovoltaic and Electronics Experiments (APEX) satellites. The TSC performed orbit reconstruction and prediction on the STEP-0 and APEX vehicles using GPS receiver navigation solution data, SGLS RAER data, and SGLS anglesonly (azimuth and elevation) data. For the STEP-O case, the navigation solution based orbits proved to be more accurate than SGLS RAER based orbits. For the APEX case, navigation solution based orbits proved to be less accurate than SGLS RAER based orbits for orbit prediction, and results for orbit reconstruction were inconclusive due to the lack of a precise truth orbit. After evaluating several different GPS data processing methods, the TSC concluded that using GPS navigation solution data is a viable alternative to using SGLS RAER data.
Spacecraft Orbit Determination with B Spline Approximation Method
NASA Astrophysics Data System (ADS)
Song, Y. Z.; Huang, Y.; Hu, X. G.; Li, P. J.; Cao, J. F.
2013-07-01
It is known that the dynamical orbit determination is the most common way to get the precise orbit of spacecraft. However, it is hard to describe the precise orbit of spacecraft sometimes. In order to solve this problem, the technique of the orbit determination with the B spline approximation method based on the theory of function approximation is presented in this article. Several simulation cases of the orbit determination including LEO (Low Earth Orbit), MEO (Medium Earth Orbit), and HEO (Highly Eccentric Orbit) satellites are performed, and it is shown that the accuracy of this method is reliable and stable.The approach can be performed in the conventional celestial coordinate system and conventional terrestrial coordinate system.The spacecraft's position and velocity can be calculated directly with the B spline approximation method, which means that it is unnecessary to integrate the dynamics equations and variational equations. In that case, it makes the calculation amount of orbit determination reduce substantially relative to the dynamical orbit determination method. The technique not only has a certain theoretical significance, but also can be as a conventional algorithm in the spacecraft orbit determination.
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.
GRAS NRT Precise Orbit Determination: Operational Experience
NASA Technical Reports Server (NTRS)
MartinezFadrique, Francisco M.; Mate, Alberto Agueda; Rodriquez-Portugal, Francisco Sancho
2007-01-01
EUMETSAT launched the meteorological satellite MetOp-A in October 2006; it is the first of the three satellites that constitute the EUMETSAT Polar System (EPS) space segment. This satellite carries a challenging and innovative instrument, the GNSS Receiver for Atmospheric Sounding (GRAS). The goal of the GRAS instrument is to support the production of atmospheric profiles of temperature and humidity with high accuracy, in an operational context, based on the bending of the GPS signals traversing the atmosphere during the so-called occultation periods. One of the key aspects associated to the data processing of the GRAS instrument is the necessity to describe the satellite motion and GPS receiver clock behaviour with high accuracy and within very strict timeliness limitations. In addition to these severe requirements, the GRAS Product Processing Facility (PPF) must be integrated in the EPS core ground segment, which introduces additional complexity from the data integration and operational procedure points of view. This paper sets out the rationale for algorithm selection and the conclusions from operational experience. It describes in detail the rationale and conclusions derived from the selection and implementation of the algorithms leading to the final orbit determination requirements (0.1 mm/s in velocity and 1 ns in receiver clock error at 1 Hz). Then it describes the operational approach and extracts the ideas and conclusions derived from the operational experience.
Ulysses orbit determination at high declinations
NASA Technical Reports Server (NTRS)
Mcelrath, Timothy P.; Lewis, George D.
1995-01-01
The trajectory of the Ulysses spacecraft caused its geocentric declination to exceed 60 deg South for over two months during the Fall of 1994, permitting continuous tracking from a single site. During this time, spacecraft operations constraints allowed only Doppler tracking data to be collected, and imposed a high radial acceleration uncertainty on the orbit determination process. The unusual aspects of this situation have motivated a re-examination of the Hamilton-Melbourne results, which have been used before to estimate the information content of Doppler tracking for trajectories closer to the ecliptic. The addition of an acceleration term to this equation is found to significantly increase the declination uncertainty for symmetric passes. In addition, a simple means is described to transform the symmetric results when the tracking pass is non-symmetric. The analytical results are then compared against numerical studies of this tracking geometry and found to be in good agreement for the angular uncertainties. The results of this analysis are applicable to the Near Earth Asteroid Rendezvous (NEAR) mission and to any other missions with high declination trajectories, as well as to missions using short tracking passes and/or one-way Doppler data.
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.
MicroGPS for Low-Cost Orbit Determination
NASA Astrophysics Data System (ADS)
Wu, S. C.; Bertiger, W. I.; Kuang, D.; Lichten, S. M.; Nandi, S.; Romans, L. J.; Srinivasan, J. M.
1997-07-01
This article presents a new technology for satellite orbit determination using a simple Global Positioning System (GPS) receiver (microGPS) with ultra-low cost, power, and mass. The capability of low-cost orbit determination with microGPS for a low Earth-orbiting satellite, Student Nitric Oxide Explorer (SNOE), is demonstrated using actual GPS data from the GPS/Meteorology (MET) satellite. The measurements acquired by the microGPS receiver will be snapshots of carrier Doppler and ambiguous pseudorange. Among the challenges in orbit determination are the resolution of the pseudorange ambiguity; the estimation of the measurement time tag drift, which effects the in-track orbit position solution; and the convergence of the orbit solution from a cold start with essentially no knowledge of the orbit. The effects of data gaps and Doppler data quality are investigated. An efficient data acquisition scenario for SNOE is derived.
Low-Earth Orbit Determination from Gravity Gradient Measurements
NASA Astrophysics Data System (ADS)
Sun, Xiucong; Chen, Pei; Macabiau, Christophe; Han, Chao
2016-06-01
An innovative orbit determination method which makes use of gravity gradients for Low-Earth-Orbiting satellites is proposed. The measurement principle of gravity gradiometry is briefly reviewed and the sources of measurement error are analyzed. An adaptive hybrid least squares batch filter based on linearization of the orbital equation and unscented transformation of the measurement equation is developed to estimate the orbital states and the measurement biases. The algorithm is tested with the actual flight data from the European Space Agency's Gravity field and steady-state Ocean Circulation Explorer (GOCE). The orbit determination results are compared with the GPS-derived orbits. The radial and cross-track position errors are on the order of tens of meters, whereas the along-track position error is over one order of magnitude larger. The gravity gradient based orbit determination method is promising for potential use in GPS-denied spacecraft navigation.
Spacecraft Orbit Determination with The B-spline Approximation Method
NASA Astrophysics Data System (ADS)
Song, Ye-zhi; Huang, Yong; Hu, Xiao-gong; Li, Pei-jia; Cao, Jian-feng
2014-04-01
It is known that the dynamical orbit determination is the most common way to get the precise orbits of spacecraft. However, it is hard to build up the precise dynamical model of spacecraft sometimes. In order to solve this problem, the technique of the orbit determination with the B-spline approximation method based on the theory of function approximation is presented in this article. In order to verify the effectiveness of this method, simulative orbit determinations in the cases of LEO (Low Earth Orbit), MEO (Medium Earth Orbit), and HEO (Highly Eccentric Orbit) satellites are performed, and it is shown that this method has a reliable accuracy and stable solution. The approach can be performed in both the conventional celestial coordinate system and the conventional terrestrial coordinate system. The spacecraft's position and velocity can be calculated directly with the B-spline approximation method, it needs not to integrate the dynamical equations, nor to calculate the state transfer matrix, thus the burden of calculations in the orbit determination is reduced substantially relative to the dynamical orbit determination method. The technique not only has a certain theoretical significance, but also can serve as a conventional algorithm in the spacecraft orbit determination.
Contribution Analysis of BDS/GPS Combined Orbit Determination
NASA Astrophysics Data System (ADS)
Zhang, Qin
2016-07-01
BeiDou Navigation Satellite System (BDS) does not have the ability of global navigation and positioning currently. The whole tracking observation of satellite orbit and the geometry of reference station are not perfect. These situations influence the accuracy of satellite orbit determination. Based on the theory and method of dynamic orbit determination, the analytical contribution of multi-GNSS combined orbit determination to the solution precision of parameters was derived. And using the measured data, the statistical contribution of BDS/GPS combined orbit determination to the solution precision of orbit and clock error was analyzed. The results show that the contribution of combined orbit determination to the solution precision of the common parameters between different systems was significant. The solution precisions of the orbit and clock error were significantly improved except GEO satellites. The statistical contribution of BDS/GPS combined orbit determination to the precision of BDS satellite orbit, the RMS of BDS satellite clock error and the RMS of receiver clock error were 36.21%, 26.88% and 20.88% respectively. Especially, the contribution to the clock error of receivers which were in the area with few visible satellites was particularly significant. And the statistical contribution was 45.95%.
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.
NASA Astrophysics Data System (ADS)
Culp, Robert D.; Mackison, Don; Fu, Ho-Ling
This paper deals with an advanced Kalman filter application to orbit determination from satellite tracking data. Modern control theory is used to set up an optimal Kalman gain for the estimation problem and to estimate its errors out of the system outputs. The classical orbit determination techniques have been used over the years for the evaluation of data analysis. A recent study was conducted to find the initial state values by modern orbit determination with Kalman gain. An original algorithm introduced by Born et al. (1986) has been applied to the spacecraft and earth satellite orbit determination for several years. The determination of the desired process and special Kalman gain for the best estimator include three kinds of computational algorithms: Batch, Sequential, and Extended Sequential processors. The model is based on a minimum variance using estimation and prediction techniques. Moreover, the estimation and computational algorithms have been modified in the UNIX system simulating to the TOPEX satellite orbit data process.
Definitive orbit determination for the HEAO-2 spacecraft
NASA Astrophysics Data System (ADS)
Smith, R. L.; Mallick, M. K.
1984-08-01
Precise ephemerides for the High Energy Astronomy Observatory-2 (HEAO-2) were computed to assist in the Charles Stark Draper Laboratory development of an onboard orbit determination technique. Weighted least-squares, batch orbit solutions were calculated using a high-precision earth gravity model and an approximate model for intermittent spacecraft thrusting. With these improvements, orbit solution consistencies at the 50- to 100-meter level were attained using the Goddard Trajectory Determination System and NASA S-band tracking data.
Definitive orbit determination for the HEAO-2 spacecraft
NASA Technical Reports Server (NTRS)
Smith, R. L.; Mallick, M. K.
1984-01-01
Precise ephemerides for the High Energy Astronomy Observatory-2 (HEAO-2) were computed to assist in the Charles Stark Draper Laboratory development of an onboard orbit determination technique. Weighted least-squares, batch orbit solutions were calculated using a high-precision earth gravity model and an approximate model for intermittent spacecraft thrusting. With these improvements, orbit solution consistencies at the 50- to 100-meter level were attained using the Goddard Trajectory Determination System and NASA S-band tracking data.
Real-time shipboard orbit determination using Kalman filtering techniques
NASA Technical Reports Server (NTRS)
Brammer, R. F.
1974-01-01
The real-time tracking and orbit determination program used on board the NASA tracking ship, the USNS Vanguard, is described in this paper. The computer program uses a variety of filtering algorithms, including an extended Kalman filter, to derive real-time orbit determinations (position-velocity state vectors) from shipboard tracking and navigation data. Results from Apollo missions are given to show that orbital parameters can be estimated quickly and accurately using these methods.
Phenomenological Determination of the Orbital Angular Momentum
Ramsey, Gordon P.
2009-08-04
Measurements involving the gluon spin, {delta}G(x, t) and the corresponding asymmetry, A(x,t) = {delta}G(x,t)/G(x,t) play an important role in quantitative understanding of proton structure. We have modeled the asymmetry perturbatively and calculated model corrections to obtain information about non-perturbative spin-orbit effects. These models are consistent with existing COMPASS and HERMES data on the gluon asymmetry. The J{sub z} = (1/2) sum rule is used to generate values of orbital angular momentum at LO and NLO. For models consistent with data, the orbital angular momentum is small. Our studies specify accuracy that future measurements should achieve to constrain theoretical models for nucleon structure.
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.
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.
TOPEX/Poseidon precision orbit determination production and expert system
NASA Technical Reports Server (NTRS)
Putney, Barbara; Zelensky, Nikita; Klosko, Steven
1993-01-01
TOPEX/Poseidon (T/P) is a joint mission between NASA and the Centre National d'Etudes Spatiales (CNES), the French Space Agency. The TOPEX/Poseidon Precision Orbit Determination Production System (PODPS) was developed at Goddard Space Flight Center (NASA/GSFC) to produce the absolute orbital reference required to support the fundamental ocean science goals of this satellite altimeter mission within NASA. The orbital trajectory for T/P is required to have a RMS accuracy of 13 centimeters in its radial component. This requirement is based on the effective use of the satellite altimetry for the isolation of absolute long-wavelength ocean topography important for monitoring global changes in the ocean circulation system. This orbit modeling requirement is at an unprecedented accuracy level for this type of satellite. In order to routinely produce and evaluate these orbits, GSFC has developed a production and supporting expert system. The PODPS is a menu driven system allowing routine importation and processing of tracking data for orbit determination, and an evaluation of the quality of the orbit so produced through a progressive series of tests. Phase 1 of the expert system grades the orbit and displays test results. Later phases undergoing implementation, will prescribe corrective actions when unsatisfactory results are seen. This paper describes the design and implementation of this orbit determination production system and the basis for its orbit accuracy assessment within the expert system.
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.
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.
Strategies for high-precision Global Positioning System orbit determination
NASA Technical Reports Server (NTRS)
Lichten, Stephen M.; Border, James S.
1987-01-01
Various strategies for the high-precision orbit determination of the GPS satellites are explored using data from the 1985 GPS field test. Several refinements to the orbit determination strategies were found to be crucial for achieving high levels of repeatability and accuracy. These include the fine tuning of the GPS solar radiation coefficients and the ground station zenith tropospheric delays. Multiday arcs of 3-6 days provided better orbits and baselines than the 8-hr arcs from single-day passes. Highest-quality orbits and baselines were obtained with combined carrier phase and pseudorange solutions.
An orbit determination from debris impacts on measurement satellites
NASA Astrophysics Data System (ADS)
Fujita, Koki; Tasaki, Mitsuhiko; Furumoto, Masahiro; Hanada, Toshiya
2016-01-01
This work proposes a method to determine orbital plane of a micron-sized space debris cloud utilizing their impacts on measurement satellites. Given that debris impacts occur on a line of intersection between debris and satellites orbital planes, a couple of debris orbital parameters, right ascension of the ascending node, inclination, and nodal regression rate can be determined by impact times and locations measured from more than two satellites in different earth orbits. This paper proves that unique solution for the debris orbital parameters is obtained from the measurement data, and derives a computational scheme to estimate them. The effectiveness of the proposed scheme is finally demonstrated by a simulation test, in which measurement data are obtained from a numerical simulation considering realistic debris' and satellites' orbits.
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
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
NASA Astrophysics Data System (ADS)
Svoren, J.; Neslusan, L.; Porubcan, V.
1994-08-01
All known parent bodies of meteor showers belong to bodies moving in high-eccentricity orbits (e => 0.5). Recently, asteroids in low-eccentricity orbits (e < 0.5) approaching the Earth's orbit, were suggested as another population of possible parent bodies of meteor streams. This paper deals with the problem of calculation of meteor radiants connected with the bodies in low-eccentricity orbits from the point of view of optimal results depending on the method applied. The paper is a continuation of our previous analysis of high-eccentricity orbits (Svoren, J., Neslusan, L., Porubcan, V.: 1993, Contrib. Astron. Obs. Skalnate Pleso 23, 23). Some additional methods resulting from mathematical modelling are presented and discussed together with Porter's, Steel-Baggaley's and Hasegawa's methods. In order to be able to compare how suitable the application of the individual radiant determination methods is, it is necessary to determine the accuracy with which they approximate real meteor orbits. To verify the accuracy with which the orbit of a meteoroid with at least one node at 1 AU fits the original orbit of the parent body, the Southworth-Hawkins D-criterion (Southworth, R.B., Hawkins, G.S.: 1963, Smithson. Contr. Astrophys. 7, 261) was applied. D <= 0.1 indicates a very good fit of orbits, 0.1 < D <= 0.2 is considered for a good fit and D > 0.2 means that the fit is rather poor and the change of orbit unrealistic. The optimal method, i.e. the one which results in the smallest D values for the population of low-eccentricity orbits, is that of adjusting the orbit by varying both the eccentricity and perihelion distance. A comparison of theoretical radiants obtained by various methods was made for typical representatives from each group of the NEA (near-Earth asteroids) objects.
NASA Astrophysics Data System (ADS)
Löcher, Anno; Kusche, Jürgen
2014-05-01
The Lunar Reconnaissance Orbiter (LRO) launched in 2009 by the National Aeronautics and Space Administration (NASA) still orbits the Moon in a polar orbit at an altitude of 50 kilometers and below. Its main objective is the detailed exploration of the Moon's surface by means of the Lunar Orbiter Laser Altimeter (LOLA) and three high resolution cameras bundled in the Lunar Reconnaissance Orbiter Camera (LROC) unit. Referring these observations to a Moon-fixed reference frame requires the computation of highly accurate and consistent orbits. For this task only Earth-based observations are available, primarily radiometric tracking data from stations in the United States, Australia and Europe. In addition, LRO is prepared for one-way laser measurements from specially adapted sites. Currently, 10 laser stations participate more or less regularly in this experiment. For operational reasons, the official LRO orbits from NASA only include radiometric data so far. In this presentation, we investigate the benefit of the laser ranging data by feeding both types of observations in an integrated orbit determination process. All computations are performed by an in-house software development based on a dynamical approach improving orbit and force parameters in an iterative way. Special attention is paid to the determination of bias parameters, in particular of timing biases between radio and laser stations and the drift and aging of the LRO spacecraft clock. The solutions from the combined data set will be compared to radio- and laser-only orbits as well as to the NASA orbits. Further results will show how recent gravity field models from the GRAIL mission can improve the accuracy of the LRO orbits.
NASA Astrophysics Data System (ADS)
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.
Contributions of Satellite Laser Ranging to the Precise Orbit Determination of Low Earth Orbiters
NASA Astrophysics Data System (ADS)
Wirnsberger, H.; Krauss, S.; Baur, O.
2014-11-01
Space-based monitoring and modeling of the system Earth requires precise knowledge of the orbits of artificial satellites. In this framework, since decades Satellite Laser Ranging (SLR) contributes with high measurement accuracy and robust tracking data to precise orbit determination. One essential role of SLR tracking is the external validation of orbit solutions derived from Global Navigation Satellite Systems (GNSS), such as the Global Positioning System (GPS). This valuable task of external validation is performed by the comparison of computed ranges based on orbit solutions and unambiguous SLR tracking data (observed ranges). Apart from validation, extension of the existing SLR network by passive antennas in combination with multistatic observations provides improvements in orbit determination processes with the background of sparse tracking data. Conceptually, these multistatic observations refer to the tracking of spacecraft from an active SLR-station and the detection of the diffuse reflected photons from the spacecraft at one or more passive stations.
Dependence of Orbit Determination Accuracy on the Observer Position
NASA Astrophysics Data System (ADS)
Vananti, Alessandro; Schildknecht, Thomas
2013-08-01
The Astronomical Institute of the University of Bern (AIUB) is conducting several search campaigns for space debris in Geostationary (GEO) and Medium Earth Orbits (MEO). Usually, to improve the quality of the determined orbits for newly discovered objects, follow-up observations are conducted. The latter take place at different times during the discovery night or in subsequent nights. The time interval between the observations plays an important role in the accuracy of the calculated orbits. Another essential parameter to consider is the position of the observer at the observation time. In this paper, the accuracy of the orbit determination with respect to the position of the observer is analyzed. The same observing site at varying epochs or multiple site locations involve different distances from the target object and a different observing angle with respect to its orbital plane and trajectory. The formal error in the orbit determination process is, among other dependencies, a function of the latter parameters. The analysis of this dependence is important to choose the appropriate observation strategy. One of the main questions that arises is e.g. whether observing the same object from different stations results in better determined orbits and, if yes, how big is the improvement. Another question is e.g. whether the observation from multiple sites needs to be simultaneous or not for a better orbit accuracy.
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)
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.
50 CFR 296.9 - Initial determination.
Code of Federal Regulations, 2011 CFR
2011-10-01
... 50 Wildlife and Fisheries 9 2011-10-01 2011-10-01 false Initial determination. 296.9 Section 296.9 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE CONTINENTAL SHELF FISHERMEN'S CONTINGENCY FUND § 296.9 Initial determination....
50 CFR 296.9 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 50 Wildlife and Fisheries 7 2010-10-01 2010-10-01 false Initial determination. 296.9 Section 296.9 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE CONTINENTAL SHELF FISHERMEN'S CONTINGENCY FUND § 296.9 Initial determination....
Determination of AES Orbit Elements Using Mixed Data
NASA Astrophysics Data System (ADS)
Kolesnik, S. Ja.; Strakhova, S. L.
An algorithm is worked out and a program is compiled for a determination of AES (artificial Earth satellite) orbit elements using both goniometrical and range-finder observations of different precision. The observations of one or several passages carried out from one or several stations can be used. A number of observational stations and a number of observations are not limited in principle. When solving this task the AES ephemerides on the moments of observations are calculated for different sets of orbit elements. A parameter F is considered which is a function of orbit elements. The parameter presents a square-mean deviation of AES ephemeris position on the moments {J;} from its observed one. The determination of real orbit elements comes to minimizing of parameter F by orbit elements using a method of deformed polyhedron. When calculating the ephemeris the amendments for 2-d, 3-d, 4-th geopotential zone harmonics are considered.
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.
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.
Space Capsule Recovery Orbit Determination System and Performance
NASA Astrophysics Data System (ADS)
Vighnesam, N. V.; Sonney, A.; Soni, P. K.
2008-08-01
Space Capsule Recovery (SRE), a small satellite, completely recoverable capsule was launched by the Polar Satellite Launch Vehicle (PSLV-C7) from the Indian spaceport Sriharikota on 10th January 2007 at 04:09UT along with Indian Remore Sensing Satellite CARTOSAT-2 and two micro satellites namely Nano- Peheunsat and Lapantubsat. The satellite was put into an almost nominal orbit of (630 X 638)km with an inclination of 97.94deg. The main objective of the SRE missions was to conduct microgravity experiment, de- orbit and recover it in Indian waters. The spacecraft was de-boosted after the payload operations in the micro- gravity environment. This was achieved in two steps. SRE was first placed from the injected circular orbit to Repetitive Elliptical Orbit (REO) and subsequently de- boosted for reentry and recovery. This paper describes the S-band based orbit determination system for SRE and its performance during different phases of the mission. Comparison of the inertial navigation system (INS) and nominal orbit with the achieved/estimated orbit was made. Orbit determination system was executed successfully through out the mission. Relatively large residues were observed in measurements during OD process due to continuous thruster activity through out the mission.
GEODYN Orbit Determination of Dawn at Vesta using Image Constraints
NASA Astrophysics Data System (ADS)
Centinello, F. J., III; Mazarico, E.; Zuber, M. T.
2012-12-01
The Dawn spacecraft has completed the orbital phase of its mapping mission of the asteroid 4 Vesta. We utilized radiometric measurements and image constraints to compute the spacecraft orbit using the GEODYN II orbit determination software. Image constraints are computed control point vectors which point from the spacecraft to landmarks observed in two images of the same region of Vesta, and are a newly developed measurement type for GEODYN. This capability was added because image constraints can provide supplemental information on the spacecraft trajectory especially in a weak gravity environment. Due to the geometric nature of image constraints, they can reduce the orbital errors in the along- and cross-track directions, which have typically carried higher uncertainty in previous interplanetary missions. Image constraints are also useful during times of absence of radiometric tracking data. Improvements to orbit determination can provide improved gravity field estimation and knowledge of the interior structure of Vesta. The NASA Deep Space Network (DSN) provides X-band tracking measurements for Dawn. Radiometric and image constraints were processed for the High Altitude Mapping Orbit (HAMO) I and II, and the Low Altitude Mapping Orbit (LAMO), from 23 Sept 2011 to 26 July 2012. The spacecraft altitude was roughly 685 km during HAMO and 200 km during LAMO. Doppler and range residual RMS were under 1 mm/s and 10 m, respectively. Improvement in orbital knowledge from image constraints is typically greatest in the cross-track direction and in our analysis these residuals were typically better than 500 m.
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2016-03-01
We present results for Precise Orbit Determination (POD) of the Lunar Reconnaissance Orbiter (LRO) based on two-way Doppler range-rates over a time span of ~13 months (January 3, 2011 to February 9, 2012). Different orbital arc lengths and various sets of empirical parameters were tested to seek optimal parametrization. An overlap analysis covering three months of Doppler data shows that the most precise orbits are obtained using an arc length of 2.5 days and estimating arc-wise constant empirical accelerations in along track direction. The overlap analysis over the entire investigated time span of 13 months indicates an orbital precision of 13.79 m, 14.17 m, and 1.28 m in along track, cross track, and radial direction, respectively, with 21.32 m in total position. We compare our orbits to the official science orbits released by the US National Aeronautics and Space Administration (NASA). The differences amount to 9.50 m, 6.98 m, and 1.50 m in along track, cross track, and radial direction, respectively, as well as 12.71 m in total position. Based on the reconstructed LRO orbits, we estimated lunar gravity field coefficients up to spherical harmonic degree and order 60. The results are compared to gravity field solutions derived from data collected by other lunar missions.
Precise orbit determination of the Lunar Reconnaissance Orbiter and first gravity field results
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2014-05-01
The Lunar Reconnaissance Orbiter (LRO) was launched in 2009 and is expected to orbit the Moon until the end of 2014. Among other instruments, LRO has a highly precise altimeter on board demanding an orbit accuracy of one meter in the radial component. Precise orbit determination (POD) is achieved with radiometric observations (Doppler range rates, ranges) on the one hand, and optical laser ranges on the other hand. LRO is the first satellite at a distance of approximately 360 000 to 400 000 km from the Earth that is routinely tracked with optical laser ranges. This measurement type was introduced to achieve orbits of higher precision than it would be possible with radiometric observations only. In this contribution we investigate the strength of each measurement type (radiometric range rates, radiometric ranges, optical laser ranges) based on single-technique orbit estimation. In a next step all measurement types are combined in a joined analysis. In addition to POD results, preliminary gravity field coefficients are presented being a subsequent product of the orbit determination process. POD and gravity field estimation was accomplished with the NASA/GSFC software packages GEODYN and SOLVE.
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.
Bayesian statistical approach to binary asteroid orbit determination
NASA Astrophysics Data System (ADS)
Kovalenko, Irina D.; Stoica, Radu S.; Emelyanov, N. V.; Doressoundiram, A.; Hestroffer, D.
2016-01-01
The problem of binary asteroids orbit determination is of particular interest, given knowledge of the orbit is the best way to derive the mass of the system. Orbit determination from observed points is a classic problem of celestial mechanics. However, in the case of binary asteroids, particularly with a small number of observations, the solution is not evident to derive. In the case of resolved binaries the problem consists in the determination of the relative orbit from observed relative positions of a secondary asteroid with respect to the primary. In this work, the problem is investigated as a statistical inverse problem. Within this context, we propose a method based on Bayesian modelling together with a global optimisation procedure that is based on the simulated annealing algorithm.
Orbit and attitude determination using artificial satellite imagery
NASA Astrophysics Data System (ADS)
Kawamura, S.; Nishida, S.; Nishimura, T.
1980-09-01
Simultaneous determination of orbital as well as attitude parameters of geostationary satellites is proposed in this paper. For this purpose, landmarks contained in the images of the earth taken by such satellites are utilized and attention is focussed on the accuracy of estimates of orbital parameters attained by this method, thus extracting effective informations on the location of the satellite contained in the images. The technology and algorithm for the simultaneous determination of orbit and attitude are actually applied to the geostationary meteorological satellite 'HIMAWARI' of Japan and the experimental results are presented. The precision of orbit determination using landmarks is less than 9 km, and it will offer useful informations to the image processing. The proposed method will be also useful at the emergency when a hazard takes place in the ranging station or ranging devices.
Robust Orbit Determination and Classification: A Learning Theoretic Approach
NASA Astrophysics Data System (ADS)
Sharma, S.; Cutler, J. W.
2015-11-01
Orbit determination involves estimation of a non-linear mapping from feature vectors associated with the position of the spacecraft to its orbital parameters. The de facto standard in orbit determination in real-world scenarios for spacecraft has been linearized estimators such as the extended Kalman filter. Such an estimator, while very accurate and convergent over its linear region, is hard to generalize over arbitrary gravitational potentials and diverse sets of measurements. It is also challenging to perform exact mathematical characterizations of the Kalman filter performance over such general systems. Here we present a new approach to orbit determination as a learning problem involving distribution regression and, also, for the multiple-spacecraft scenario, a transfer learning system for classification of feature vectors associated with spacecraft, and provide some associated analysis of such systems.
Precise orbit determination based on raw GPS measurements
NASA Astrophysics Data System (ADS)
Zehentner, Norbert; Mayer-Gürr, Torsten
2016-03-01
Precise orbit determination is an essential part of the most scientific satellite missions. Highly accurate knowledge of the satellite position is used to geolocate measurements of the onboard sensors. For applications in the field of gravity field research, the position itself can be used as observation. In this context, kinematic orbits of low earth orbiters (LEO) are widely used, because they do not include a priori information about the gravity field. The limiting factor for the achievable accuracy of the gravity field through LEO positions is the orbit accuracy. We make use of raw global positioning system (GPS) observations to estimate the kinematic satellite positions. The method is based on the principles of precise point positioning. Systematic influences are reduced by modeling and correcting for all known error sources. Remaining effects such as the ionospheric influence on the signal propagation are either unknown or not known to a sufficient level of accuracy. These effects are modeled as unknown parameters in the estimation process. The redundancy in the adjustment is reduced; however, an improvement in orbit accuracy leads to a better gravity field estimation. This paper describes our orbit determination approach and its mathematical background. Some examples of real data applications highlight the feasibility of the orbit determination method based on raw GPS measurements. Its suitability for gravity field estimation is presented in a second step.
NASA Technical Reports Server (NTRS)
Lichten, S. M.; Estefan, J. A.
1990-01-01
Orbit covariance analyses pertaining to the Japanese VLBI Space Observatory Program (VSOP) MUSES-B satellite and to the International VLBI Satellite are presented. It is determined that a combination of Doppler and GPS measurements can provide the orbit accuracy required to support advanced radio interferometric experiments. For the VSOP, the required orbit accuracy of 130 m is easily met with two-way Doppler as the primary type of data; the 0.4 cm/s VSOP velocity requirement is also feasible provided that precise ground calibrations of tropospheric delays and station coordinates are available. It is concluded that combining the data from a VSOP GPS flight instrument with the ground GPS and two-way Doppler data will significantly enhance orbit determination accuracy in position and velocity.
Orbit Determination and Navigation Software Testing for the Mars Reconnaissance Orbiter
NASA Technical Reports Server (NTRS)
Pini, Alex
2011-01-01
During the extended science phase of the Mars Reconnaissance Orbiter's lifecycle, the operational duties pertaining to navigation primarily involve orbit determination. The orbit determination process utilizes radiometric tracking data and is used for the prediction and reconstruction of MRO's trajectories. Predictions are done twice per week for ephemeris updates on-board the spacecraft and for planning purposes. Orbit Trim Maneuvers (OTM-s) are also designed using the predicted trajectory. Reconstructions, which incorporate a batch estimator, provide precise information about the spacecraft state to be synchronized with scientific measurements. These tasks were conducted regularly to validate the results obtained by the MRO Navigation Team. Additionally, the team is in the process of converting to newer versions of the navigation software and operating system. The capability to model multiple densities in the Martian atmosphere is also being implemented. However, testing outputs among these different configurations was necessary to ensure compliance to a satisfactory degree.
Precision orbit determination using TOPEX/Poseidon TDRSS observations
NASA Technical Reports Server (NTRS)
Teles, Jerome; Putney, B.; Phelps, J.; Mccarthy, J.; Eddy, W.; Klosko, S.
1993-01-01
The TOPEX/Poseidon (T/P) Mission carries a variety of packages to support experimental, precision and operational orbit determination. Included are a GPS transponder, laser retro-reflectors, the French-developed Doppler Orbitography and Radiopositioning Integrated by Satellite (DORIS) Doppler tracking system and a Tracking Data Relay Satellite System (TDRSS) transponder. Presently, TDRSS tracking is used for operational orbit support and is processed with force and measurement modeling consistent with this purpose. However, the low noise and extensive geographical coverage of the TDRSS/TOPEX data allows an assessment of TDRSS Precision Orbit Determination (POD) capabilities by comparison to the T/P precision orbit determination. The Geodynamics (GEODYN) Orbit Determination System is used to process laser and DORIS data to produce the precision orbits for the T/P Project. GEODYN has been modified recently to support the TDRSS observations. TDRSS data analysis can now benefit from the extensive force modeling and reference frame stability needed to meet the orbit determination (OD) goals of the T/P Mission. This analysis has concentrated on the strongest of the TDRSS measurement types, its two-way average range rate. Both the TDRSS and T/P orbits have been assessed in combination with the global satellite laser ranging (SLR) data and by themselves. These results indicate that significant improvement in the TDRSS ephemerides is obtained when the T/P orbit is well determined by SLR, and the TDRSS/TOPEX Doppler link is used to position TDRSS. Meter-level TDRSS positioning uncertainty is achieved using this approach. When the TDRSS orbit location is provided by this approach, the two-way range rate from a single TDRSS (i.e. West only) can provide T/P orbits with sub-meter radial accuracies and two meter RMS total position agreement with SLR defined orbits. These preliminary results indicate improved modeling of the TDRSS measurement through the elimination of heretofore
NASA Technical Reports Server (NTRS)
Bryant, W. C., Jr.; Goad, C. C.
1973-01-01
A Tracking Data Relay Satellite System (TDRSS) made up of two earth synchronous data relay satellites is proposed for the late 1970s to aid in the tracking, or take the place of ground tracking, or near-earth orbiters. Theoretical error analysis studies were conducted to evaluate the TDRSS concept of tracking user satellites. All major factors affecting orbit determination accuracy were considered in the analysis, including tracking system and dynamic modeling errors.
NASA Astrophysics Data System (ADS)
Svoren, J.; Neslusan, L.; Porubcan, V.
1993-07-01
It is evident that there is no uniform method of calculating meteor radiants which would yield reliable results for all types of cometary orbits. In the present paper an analysis of this problem is presented, together with recommended methods for various types of orbits. Some additional methods resulting from mathematical modelling are presented and discussed together with Porter's, Steel-Baggaley's and Hasegawa's methods. In order to be able to compare how suitable the application of the individual radiant determination methods is, it is necessary to determine the accuracy with which they approximate real meteor orbits. To verify the accuracy with which the orbit of a meteoroid with at least one node at 1 AU fits the original orbit of the parent body, we applied the Southworth-Hawkins D-criterion (Southworth, R.B., Hawkins, G.S.: 1963, Smithson. Contr. Astrophys 7, 261). D<=0.1 indicates a very good fit of orbits, 0.1
Distance-based relative orbital elements determination for formation flying system
NASA Astrophysics Data System (ADS)
He, Yanchao; Xu, Ming; Chen, Xi
2016-01-01
The present paper deals with determination of relative orbital elements based only on distance between satellites in the formation flying system, which has potential application in engineering, especially suited for rapid orbit determination required missions. A geometric simplification is performed to reduce the formation configuration in three-dimensional space to a plane. Then the equivalent actual configuration deviating from its nominal design is introduced to derive a group of autonomous linear equations on the mapping between the relative orbital elements differences and distance errors. The primary linear equations-based algorithm is initially proposed to conduct the rapid and precise determination of the relative orbital elements without the complex computation, which is further improved by least-squares method with more distance measurements taken into consideration. Numerical simulations and comparisons with traditional approaches are presented to validate the effectiveness of the proposed methods. To assess the performance of the two proposed algorithms, accuracy validation and Monte Carlo simulations are implemented in the presence of noises of distance measurements and the leader's absolute orbital elements. It is demonstrated that the relative orbital elements determination accuracy of two approaches reaches more than 90% and even close to the actual values for the least-squares improved one. The proposed approaches can be alternates for relative orbit determination without assistance of additional facilities in engineering for their fairly high efficiency with accuracy and autonomy.
Orbit Determination Accuracy for Comets on Earth-Impacting Trajectories
NASA Technical Reports Server (NTRS)
Kay-Bunnell, Linda
2004-01-01
The results presented show the level of orbit determination accuracy obtainable for long-period comets discovered approximately one year before collision with Earth. Preliminary orbits are determined from simulated observations using Gauss' method. Additional measurements are incorporated to improve the solution through the use of a Kalman filter, and include non-gravitational perturbations due to outgassing. Comparisons between observatories in several different circular heliocentric orbits show that observatories in orbits with radii less than 1 AU result in increased orbit determination accuracy for short tracking durations due to increased parallax per unit time. However, an observatory at 1 AU will perform similarly if the tracking duration is increased, and accuracy is significantly improved if additional observatories are positioned at the Sun-Earth Lagrange points L3, L4, or L5. A single observatory at 1 AU capable of both optical and range measurements yields the highest orbit determination accuracy in the shortest amount of time when compared to other systems of observatories.
Evaluation of Improved Spacecraft Models for GLONASS Orbit Determination
NASA Astrophysics Data System (ADS)
Weiss, J. P.; Sibthorpe, A.; Harvey, N.; Bar-Sever, Y.; Kuang, D.
2010-12-01
High-fidelity spacecraft models become more important as orbit determination strategies achieve greater levels of precision and accuracy. In this presentation, we assess the impacts of new solar radiation pressure and attitude models on precise orbit determination (POD) for GLONASS spacecraft within JPLs GIPSY-OASIS software. A new solar radiation pressure model is developed by empirically fitting a Fourier expansion to solar pressure forces acting on the spacecraft X, Y, Z components using one year of recent orbit data. Compared to a basic “box-wing” solar pressure model, the median 24-hour orbit prediction accuracy for one month of independent test data improves by 43%. We additionally implement an updated yaw attitude model during eclipse periods. We evaluate the impacts of both models on post-processed POD solutions spanning 6-months. We consider a number of metrics such as internal orbit and clock overlaps as well as comparisons to independent solutions. Improved yaw attitude modeling reduces the dependence of these metrics on the “solar elevation” angle. The updated solar pressure model improves orbit overlap statistics by several mm in the median sense and centimeters in the max sense (1D). Orbit differences relative to the IGS combined solution are at or below the 5 cm level (1D RMS).
GRAIL Orbit Determination for the Science Phase and Extended Mission
NASA Technical Reports Server (NTRS)
Ryne, Mark; Antreasian, Peter; Broschart, Stephen; Criddle, Kevin; Gustafson, Eric; Jefferson, David; Lau, Eunice; Ying Wen, Hui; You, Tung-Han
2013-01-01
The Gravity Recovery and Interior Laboratory Mission (GRAIL) is the 11th mission of the NASA Discovery Program. Its objective is to help answer funda-mental questions about the Moon's internal structure, thermal evolution, and collisional history. GRAIL employs twin spacecraft, which fly in formation in low altitude polar orbits around the Moon. An improved global lunar gravity field is derived from high-precision range-rate measurements of the distance between the two spacecraft. The purpose of this paper is to describe the strategies used by the GRAIL Orbit Determination Team to overcome challenges posed during on-orbit operations.
NASA Astrophysics Data System (ADS)
Maier, Andrea; Baur, Oliver
2015-04-01
The Lunar Reconnaissance Orbiter (LRO), launched in 2009, is well suited for the estimation of the long wavelengths of the lunar gravity field due to its low altitude of 50 km. Further, the orbit of LRO was polar for two years providing global coverage. The satellite has been primarily tracked via S-band (mainly two-way Doppler range-rates and two-way radiometric ranges) from the dedicated station in White Sands and from the Universal Space Network (USN). Due to the onboard altimeter the orbital precision requirement in the radial direction was rigorously defined as 1m. Because simulation studies before LRO's launch showed that this precision could not be reached with S-band observations alone, it was decided to additionally track LRO via optical laser ranges. It is worthwhile to point out that LRO is the first spacecraft in interplanetary space routinely tracked with optical one-way laser ranges. Gravity field recovery from orbit perturbations is intrinsically related to precise orbit determination. This is why considerable effort was made to find the optimum settings for orbit modeling. For a time span of three months we conducted a series of orbit overlapping tests based on Doppler observations to find the optimum arc length and the optimum set of empirical parameters. The analysis of observation residuals and orbit overlap differences showed that the estimated orbits are most precise when subdividing the time span into 2.5 days and estimating one constant empirical acceleration in along track direction. These settings were then used to analyze 13 months of Doppler data to LRO. The processing of the optical one-way laser was difficult due to the involvement of two non-synchronous clocks in one measurement (one clock at the ground station and one clock onboard LRO). The NASA software GEODYN, which was used for orbit determination and parameter estimation, models the LRO clock using a drift rate (first-order term) and an aging rate (second-order term). It seems
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
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.
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
Orbit determination support of the Ocean Topography Experiment (TOPEX)/Poseidon operational orbit
NASA Technical Reports Server (NTRS)
Schanzle, A. F.; Rovnak, J. E.; Bolvin, D. T.; Doll, C. E.
1993-01-01
The Ocean Topography Experiment (TOPEX/Poseidon) mission is designed to determine the topography of the Earth's sea surface over a 3-year period, beginning shortly after launch in July 1992. TOPEX/Poseidon is a joint venture between the United States National Aeronautics and Space Administration (NASA) and the French Centre Nationale d'Etudes Spatiales. The Jet Propulsion Laboratory is NASA's TOPEX/Poseidon project center. The Tracking and Data Relay Satellite System (TDRSS) will nominally be used to support the day-to-day orbit determination aspects of the mission. Due to its extensive experience with TDRSS tracking data, the NASA Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will receive and process TDRSS observational data. To fulfill the scientific goals of the mission, it is necessary to achieve and maintain a very precise orbit. The most stringent accuracy requirements are associated with planning and evaluating orbit maneuvers, which will place the spacecraft in its mission orbit and maintain the required ground track. To determine if the FDF can meet the TOPEX/Poseidon maneuver accuracy requirements, covariance analysis was undertaken with the Orbit Determination Error Analysis System (ODEAS). The covariance analysis addressed many aspects of TOPEX/Poseidon orbit determination, including arc length, force models, and other processing options. The most recent analysis has focused on determining the size of the geopotential field necessary to meet the maneuver support requirements. Analysis was undertaken with the full 50 x 50 Goddard Earth Model (GEM) T3 field as well as smaller representations of this model.
Evaluation of semiempirical atmospheric density models for orbit determination applications
NASA Technical Reports Server (NTRS)
Cox, C. M.; Feiertag, R. J.; Oza, D. H.; Doll, C. E.
1994-01-01
This paper presents the results of an investigation of the orbit determination performance of the Jacchia-Roberts (JR), mass spectrometer incoherent scatter 1986 (MSIS-86), and drag temperature model (DTM) atmospheric density models. Evaluation of the models was performed to assess the modeling of the total atmospheric density. This study was made generic by using six spacecraft and selecting time periods of study representative of all portions of the 11-year cycle. Performance of the models was measured for multiple spacecraft, representing a selection of orbit geometries from near-equatorial to polar inclinations and altitudes from 400 kilometers to 900 kilometers. The orbit geometries represent typical low earth-orbiting spacecraft supported by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). The best available modeling and orbit determination techniques using the Goddard Trajectory Determination System (GTDS) were employed to minimize the effects of modeling errors. The latest geopotential model available during the analysis, the Goddard earth model-T3 (GEM-T3), was employed to minimize geopotential model error effects on the drag estimation. Improved-accuracy techniques identified for TOPEX/Poseidon orbit determination analysis were used to improve the Tracking and Data Relay Satellite System (TDRSS)-based orbit determination used for most of the spacecraft chosen for this analysis. This paper shows that during periods of relatively quiet solar flux and geomagnetic activity near the solar minimum, the choice of atmospheric density model used for orbit determination is relatively inconsequential. During typical solar flux conditions near the solar maximum, the differences between the JR, DTM, and MSIS-86 models begin to become apparent. Time periods of extreme solar activity, those in which the daily and 81-day mean solar flux are high and change rapidly, result in significant differences between the models. During periods of high
Unilateral Eyelid Edema as Initial Sign of Orbital Sarcoidosis.
Petrarolha, Sílvia Miguéis Picado; Rodrigues, Bruna Suda; Filho, Flávio David Haddad; Dedivitis, Rogério Aparecido; Petrarolha, Samuel Brunini; Morais, Pedro Martins Tavares Scianni
2016-01-01
Introduction. Sarcoidosis is a rare multisystemic granulomatous inflammatory disease of unknown etiology affecting the respiratory system, skin, and eyes. Sarcoidosis outside the lacrimal gland is rare. The case study concerns a patient with a final diagnosis of orbital sarcoidosis. Case Report. A 37-year-old male patient went to the ophthalmic emergency room complaining of pain in the left eye, diplopia, and decreased visual acuity. An external eye examination showed hard and cold edema of the lower eyelid, ocular motility with limitation of adduction, and discreet ipsilateral proptosis. Magnetic resonance of the orbit showed left eye proptosis and thickening and increase of soft tissues associated with heterogeneous impregnation of contrast in the infralateral region of the left eyelid. A biopsy of the lesion showed a chronic inflammatory process, with numerous compact nonnecrotizing granulomas surrounded by lamellar hyaline collagen, providing histological confirmation of sarcoidosis. Discussion. A biopsy of the orbital tumor is essential for the diagnosis of sarcoidosis, in addition to the search for systemic findings such as hilar adenopathy or parenchymal lung disease found in 90% of patients. PMID:27298746
Unilateral Eyelid Edema as Initial Sign of Orbital Sarcoidosis
Filho, Flávio David Haddad; Dedivitis, Rogério Aparecido; Petrarolha, Samuel Brunini
2016-01-01
Introduction. Sarcoidosis is a rare multisystemic granulomatous inflammatory disease of unknown etiology affecting the respiratory system, skin, and eyes. Sarcoidosis outside the lacrimal gland is rare. The case study concerns a patient with a final diagnosis of orbital sarcoidosis. Case Report. A 37-year-old male patient went to the ophthalmic emergency room complaining of pain in the left eye, diplopia, and decreased visual acuity. An external eye examination showed hard and cold edema of the lower eyelid, ocular motility with limitation of adduction, and discreet ipsilateral proptosis. Magnetic resonance of the orbit showed left eye proptosis and thickening and increase of soft tissues associated with heterogeneous impregnation of contrast in the infralateral region of the left eyelid. A biopsy of the lesion showed a chronic inflammatory process, with numerous compact nonnecrotizing granulomas surrounded by lamellar hyaline collagen, providing histological confirmation of sarcoidosis. Discussion. A biopsy of the orbital tumor is essential for the diagnosis of sarcoidosis, in addition to the search for systemic findings such as hilar adenopathy or parenchymal lung disease found in 90% of patients. PMID:27298746
Precise orbit determination of Beidou Satellites at GFZ
NASA Astrophysics Data System (ADS)
Deng, Zhiguo; Ge, Maorong; Uhlemann, Maik; Zhao, Qile
2014-05-01
In December 2012 the Signal-In-Space Interface Control Document (ICD) of the BeiDou Navigation Satellite System (BeiDou system) was published. Currently the initial BeiDou regional navigation satellite system consisting of 14 satellites was completed, and provides observation data of five Geostationary-Earth-Orbit (GEO)satellites, five Inclined-GeoSynchronous-Orbit (IGSO) satellites and four Medium-Earth-Orbit (MEO) satellites. The Helmholtz Centre Potsdam GFZ German Research Centre for Geosciences (GFZ) contributes as one of the analysis centers to the International GNSS Service (IGS) since many years. In 2012 the IGS began the "Multi GNSS EXperiment" (MGEX), which supports the new GNSS, such as Galileo, Compass, and QZSS. Based on tracking data of BeiDou-capable receivers from the MGEX and chinese BeiDou networks up to 45 global distributed stations are selected to estimate orbit and clock parameters of the GPS/BeiDou satellites. Some selected results from the combined GPS/BeiDou data processing with 10 weeks of data from 2013 are shown. The quality of the orbit and clock products are assessed by means of orbit overlap statistics, clock stabilities as well as an independent validation with SLR measurements. At the end an outlook about GFZ AC's future Multi-GNSS activities will be given.
Precise Orbit Determination of BeiDou Navigation Satellite System
NASA Astrophysics Data System (ADS)
He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald
2013-04-01
China has been developing its own independent satellite navigation system since decades. Now the COMPASS system, also known as BeiDou, is emerging and gaining more and more interest and attention in the worldwide GNSS communities. The current regional BeiDou system is ready for its operational service around the end of 2012 with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit satellites (IGSO) and four Medium Earth orbit (MEO) satellites in operation. Besides the open service with positioning accuracy of around 10m which is free to civilian users, both precise relative positioning, and precise point positioning are demonstrated as well. In order to enhance the BeiDou precise positioning service, Precise Orbit Determination (POD) which is essential of any satellite navigation system has been investigated and studied thoroughly. To further improving the orbits of different types of satellites, we study the impact of network coverage on POD data products by comparing results from tracking networks over the Chinese territory, Asian-Pacific, Asian and of global scale. Furthermore, we concentrate on the improvement of involving MEOs on the orbit quality of GEOs and IGSOs. POD with and without MEOs are undertaken and results are analyzed. Finally, integer ambiguity resolution which brings highly improvement on orbits and positions with GPS data is also carried out and its effect on POD data products is assessed and discussed in detail. Seven weeks of BeiDou data from a ground tracking network, deployed by Wuhan University is employed in this study. The test constellation includes four GEO, five IGSO and two MEO satellites in operation. The three-day solution approach is employed to enhance its strength due to the limited coverage of the tracking network and the small movement of most of the satellites. A number of tracking scenarios and processing schemas are identified and processed and overlapping orbit
Implementation of a low-cost, commercial orbit determination system
NASA Technical Reports Server (NTRS)
Corrigan, Jim
1994-01-01
Traditional satellite and launch control systems have consisted of custom solutions requiring significant development and maintenance costs. These systems have typically been designed to support specific program requirements and are expensive to modify and augment after delivery. The expanding role of space in today's marketplace combined with the increased sophistication and capabilities of modern satellites has created a need for more efficient, lower cost solutions to complete command and control systems. Recent technical advances have resulted in commercial-off-the-shelf products which greatly reduce the complete life-cycle costs associated with satellite launch and control system procurements. System integrators and spacecraft operators have, however, been slow to integrate these commercial based solutions into a comprehensive command and control system. This is due, in part, to a resistance to change and the fact that many available products are unable to effectively communicate with other commercial products. The United States Air Force, responsible for the health and safety of over 84 satellites via its Air Force Satellite Control Network (AFSCN), has embarked on an initiative to prove that commercial products can be used effectively to form a comprehensive command and control system. The initial version of this system is being installed at the Air Force's Center for Research Support (CERES) located at the National Test Facility in Colorado Springs, Colorado. The first stage of this initiative involved the identification of commercial products capable of satisfying each functional element of a command and control system. A significant requirement in this product selection criteria was flexibility and ability to integrate with other available commercial products. This paper discusses the functions and capabilities of the product selected to provide orbit determination functions for this comprehensive command and control system.
Implementation of a low-cost, commercial orbit determination system
NASA Astrophysics Data System (ADS)
Corrigan, Jim
1994-11-01
Traditional satellite and launch control systems have consisted of custom solutions requiring significant development and maintenance costs. These systems have typically been designed to support specific program requirements and are expensive to modify and augment after delivery. The expanding role of space in today's marketplace combined with the increased sophistication and capabilities of modern satellites has created a need for more efficient, lower cost solutions to complete command and control systems. Recent technical advances have resulted in commercial-off-the-shelf products which greatly reduce the complete life-cycle costs associated with satellite launch and control system procurements. System integrators and spacecraft operators have, however, been slow to integrate these commercial based solutions into a comprehensive command and control system. This is due, in part, to a resistance to change and the fact that many available products are unable to effectively communicate with other commercial products. The United States Air Force, responsible for the health and safety of over 84 satellites via its Air Force Satellite Control Network (AFSCN), has embarked on an initiative to prove that commercial products can be used effectively to form a comprehensive command and control system. The initial version of this system is being installed at the Air Force's Center for Research Support (CERES) located at the National Test Facility in Colorado Springs, Colorado. The first stage of this initiative involved the identification of commercial products capable of satisfying each functional element of a command and control system. A significant requirement in this product selection criteria was flexibility and ability to integrate with other available commercial products. This paper discusses the functions and capabilities of the product selected to provide orbit determination functions for this comprehensive command and control system.
Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation
NASA 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
Precise GPS orbit determination results from 1985 field tests
NASA Technical Reports Server (NTRS)
Lichten, S. M.; Border, J. S.; Wu, S.-C.; Williams, B. G.; Yunck, T. P.
1986-01-01
Data from three different receiver types have been used to obtain precise orbits for the satellites of the Global Positioning System (GPS). The data were collected during the 1985 March-April GPS experiment to test and validate GPS techniques for precision orbit determination and geodesy. A new software package developed at the Jet Propulsion Laboratory (JPL), GIPSY (GPS Inferred Positioning SYstem), was used to process the data. To assess orbit accuracy, solutions are compared using integrated doppler data from various different receiver types, different fiducial sites, and independent data arcs, including one spanning six days. From these intercomparisons, orbit accuracy for a well-tracked GPS satellite of three meters in altitude and about five meters in each of down and cross-track components are inferred.
Copernicus POD Service: Orbit Determination of the Sentinel Satellites
NASA Astrophysics Data System (ADS)
Peter, Heike; Fernández, Jaime; Ayuga, Francisco; Féménias, Pierre
2016-04-01
The Copernicus POD (Precise Orbit Determination) Service is part of the Copernicus Processing Data Ground Segment (PDGS) of the Sentinel-1, -2 and -3 missions. A GMV-led consortium is operating the Copernicus POD Service being in charge of generating precise orbital products and auxiliary data files for their use as part of the processing chains of the respective Sentinel PDGS. Sentinel-1A was launched in April 2014 while Sentinel-2A was on June 2015 and both are routinely operated since then. Sentinel-3A is expected to be launched in February 2016 and Sentinel-1B is planned for spring 2016. Thus the CPOD Service will be operating three to four satellites simultaneously in spring 2016. The satellites of the Sentinel-1, -2, and -3 missions are all equipped with dual frequency high precision GPS receivers delivering the main observables for POD. Sentinel-3 satellites will additionally be equipped with a laser retro reflector for Satellite Laser Ranging and a receiver for DORIS tracking. All three types of observables (GPS, SLR and DORIS) will be used routinely for POD. The POD core of the CPOD Service is NAPEOS (Navigation Package for Earth Orbiting Satellites) the leading ESA/ESOC software for precise orbit determination. The careful selection of models and inputs is important to achieve the different but very demanding requirements in terms of orbital accuracy and timeliness for the Sentinel -1, -2 & -3 missions. The three missions require orbital products with various latencies from 30 minutes up to 20-30 days. The accuracy requirements are also different and partly very challenging, targeting 5 cm in 3D for Sentinel-1 and 2-3 cm in radial direction for Sentinel-3. Although the characteristics and the requirements are different for the three missions the same core POD setup is used to the largest extent possible. This strategy facilitates maintenance of the complex system of the CPOD Service. Updates in the dynamical modelling of the satellite orbits, e
Optimal Control for a Cooperative Rendezvous Between Two Spacecraft from Determined Orbits
NASA Astrophysics Data System (ADS)
Feng, Weiming; Han, Liping; Shi, Lei; Zhao, Di; Yang, Kun
2016-03-01
The mathematical model of a far-distance cooperative rendezvous between two spacecraft in a non-Keplerian orbit was established. Approximate global optimization was performed by a type of hybrid algorithm consisting of particle swarm optimization and differential evolution. In this process, the double-fitness function was established according to the objective function and the constraints; the double-fitness function was used to enable a better choice between the solutions obtained by the two algorithms at every iteration. In addition, the costate variables obtained were set as the initial values of the sequential quadratic programming to greatly increase the possibility of finding the approximate global optimal solution. After performing the calculations and simulations, it was concluded that the fuel required for orbiting was not influenced by the initial positions of the two spacecraft if the initial orbits of the two spacecraft were determined. However, the time consumption is strongly influenced in this situation.
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.
Evaluation of advanced geopotential models for operational orbit determination
NASA Technical Reports Server (NTRS)
Radomski, M. S.; Davis, B. E.; Samii, M. V.; Engel, C. J.; Doll, C. E.
1988-01-01
To meet future orbit determination accuracy requirements for different NASA projects, analyses are performed using Tracking and Data Relay Satellite System (TDRSS) tracking measurements and orbit determination improvements in areas such as the modeling of the Earth's gravitational field. Current operational requirements are satisfied using the Goddard Earth Model-9 (GEM-9) geopotential model with the harmonic expansion truncated at order and degree 21 (21-by-21). This study evaluates the performance of 36-by-36 geopotential models, such as the GEM-10B and Preliminary Goddard Solution-3117 (PGS-3117) models. The Earth Radiation Budget Satellite (ERBS) and LANDSAT-5 are the spacecraft considered in this study.
Use of the VLBI delay observable for orbit determination of Earth-orbiting VLBI satellites
NASA Technical Reports Server (NTRS)
Ulvestad, J. S.
1992-01-01
Very long-baseline interferometry (VLBI) observations using a radio telescope in Earth orbit were performed first in the 1980s. Two spacecraft dedicated to VLBI are scheduled for launch in 1995; the primary scientific goals of these missions will be astrophysical in nature. This article addresses the use of space VLBI delay data for the additional purpose of improving the orbit determination of the Earth-orbiting spacecraft. In an idealized case of quasi-simultaneous observations of three radio sources in orthogonal directions, analytical expressions are found for the instantaneous spacecraft position and its error. The typical position error is at least as large as the distance corresponding to the delay measurement accuracy but can be much greater for some geometries. A number of practical considerations, such as system noise and imperfect calibrations, set bounds on the orbit-determination accuracy realistically achievable using space VLBI delay data. These effects limit the spacecraft position accuracy to at least 35 cm (and probably 3 m or more) for the first generation of dedicated space VLBI experiments. Even a 35-cm orbital accuracy would fail to provide global VLBI astrometry as accurate as ground-only VLBI. Recommended charges in future space VLBI missions are unlikely to make space VLBI competitive with ground-only VLBI in global astrometric measurements.
Cassini orbit determination performance during the first eight orbits of the Saturn satellite tour
NASA Technical Reports Server (NTRS)
Antreasian, P. G.; Bordi, J. J.; Criddle, K. E.; Ionasescu, R.; Jacobson, R. A.; Jones, J. B.; MacKenzie, R. A.; Meek, M. C.; Pelletier, F. J.; Roth, D. C.; Roundhill, I. M.; Stauch, J.
2005-01-01
From June 2004 through July 2005, the Cassini/Huygens spacecraft has executed nine successful close-targeted encounters by three major satellites of the Saturnian system. Current results show that orbit determination has met design requirements for targeting encounters, Hugens descent, and predicting science instrument pointing for targetd satellite encounters. This paper compares actual target dispersion against, the predicte tour covariance analyses.
Analysis of orbital configurations for geocenter determination with GPS and low-Earth orbiters
NASA Astrophysics Data System (ADS)
Kuang, Da; Bar-Sever, Yoaz; Haines, Bruce
2015-05-01
We use a series of simulated scenarios to characterize the observability of geocenter location with GPS tracking data. We examine in particular the improvement realized when a GPS receiver in low Earth orbit (LEO) augments the ground network. Various orbital configurations for the LEO are considered and the observability of geocenter location based on GPS tracking is compared to that based on satellite laser ranging (SLR). The distance between a satellite and a ground tracking-site is the primary measurement, and Earth rotation plays important role in determining the geocenter location. Compared to SLR, which directly and unambiguously measures this distance, terrestrial GPS observations provide a weaker (relative) measurement for geocenter location determination. The estimation of GPS transmitter and receiver clock errors, which is equivalent to double differencing four simultaneous range measurements, removes much of this absolute distance information. We show that when ground GPS tracking data are augmented with precise measurements from a GPS receiver onboard a LEO satellite, the sensitivity of the data to geocenter location increases by more than a factor of two for Z-component. The geometric diversity underlying the varying baselines between the LEO and ground stations promotes improved global observability, and renders the GPS technique comparable to SLR in terms of information content for geocenter location determination. We assess a variety of LEO orbital configurations, including the proposed orbit for the geodetic reference antenna in space mission concept. The results suggest that a retrograde LEO with altitude near 3,000 km is favorable for geocenter determination.
GPS-Based Reduced Dynamic Orbit Determination Using Accelerometer Data
NASA Technical Reports Server (NTRS)
VanHelleputte, Tom; Visser, Pieter
2007-01-01
Currently two gravity field satellite missions, CHAMP and GRACE, are equipped with high sensitivity electrostatic accelerometers, measuring the non-conservative forces acting on the spacecraft in three orthogonal directions. During the gravity field recovery these measurements help to separate gravitational and non-gravitational contributions in the observed orbit perturbations. For precise orbit determination purposes all these missions have a dual-frequency GPS receiver on board. The reduced dynamic technique combines the dense and accurate GPS observations with physical models of the forces acting on the spacecraft, complemented by empirical accelerations, which are stochastic parameters adjusted in the orbit determination process. When the spacecraft carries an accelerometer, these measured accelerations can be used to replace the models of the non-conservative forces, such as air drag and solar radiation pressure. This approach is implemented in a batch least-squares estimator of the GPS High Precision Orbit Determination Software Tools (GHOST), developed at DLR/GSOC and DEOS. It is extensively tested with data of the CHAMP and GRACE satellites. As accelerometer observations typically can be affected by an unknown scale factor and bias in each measurement direction, they require calibration during processing. Therefore the estimated state vector is augmented with six parameters: a scale and bias factor for the three axes. In order to converge efficiently to a good solution, reasonable a priori values for the bias factor are necessary. These are calculated by combining the mean value of the accelerometer observations with the mean value of the non-conservative force models and empirical accelerations, estimated when using these models. When replacing the non-conservative force models with accelerometer observations and still estimating empirical accelerations, a good orbit precision is achieved. 100 days of GRACE B data processing results in a mean orbit fit of
The role of laser determined orbits in geodesy and geophysics
NASA Technical Reports Server (NTRS)
Kolenkiewicz, R.; Smith, D. E.; Dunn, P. J.; Torrence, M. H.; Robbins, J. W.
1991-01-01
Some of the results of orbit analysis from the NASA SLR analysis group are presented. The earth's orientation was determined for 5-day intervals to 1.9 mas for the pole and 0.09 msec for length of day. The 3d center of mass station positions was determined to 33 mm over a period of 3 months, and geodesic rates of SLR tracking sites were determined to 5 mm/yr.
GRAIL Science Data System Orbit Determination : Approach, Strategy, and Performance
NASA Technical Reports Server (NTRS)
Fahnestock, Eugene; Asmar, Sami; Park, Ryan; Strekalov, Dmitry; Yuan, Dah-Ning; Harvey, Nate; Kahan, Daniel; Konopliv, Alex; Kruizinga, Gerhard; Oudrhiri, Kamal; Paik, Meegyeong
2013-01-01
This paper details orbit determination techniques and strategies employed within each stage of the larger iterative process of preprocessing raw GRAIL data into the gravity science measurements used within gravity field solutions. Each orbit determination pass used different data, corrections to them, and/or estimation parameters. We compare performance metrics among these passes. For example, for the primary mission, the magnitude of residuals using our orbits progressed from approximately or equal to19.4 to 0.077 approximately or equal to m/s for inter-satellite range rate data and from approximately or equal to 0.4 to approximately or equal to 0.1 mm/s for Doppler data.
An intelligent interface for satellite operations: Your Orbit Determination Assistant (YODA)
NASA Technical Reports Server (NTRS)
Schur, Anne
1988-01-01
An intelligent interface is often characterized by the ability to adapt evaluation criteria as the environment and user goals change. Some factors that impact these adaptations are redefinition of task goals and, hence, user requirements; time criticality; and system status. To implement adaptations affected by these factors, a new set of capabilities must be incorporated into the human-computer interface design. These capabilities include: (1) dynamic update and removal of control states based on user inputs, (2) generation and removal of logical dependencies as change occurs, (3) uniform and smooth interfacing to numerous processes, databases, and expert systems, and (4) unobtrusive on-line assistance to users of concepts were applied and incorporated into a human-computer interface using artificial intelligence techniques to create a prototype expert system, Your Orbit Determination Assistant (YODA). YODA is a smart interface that supports, in real teime, orbit analysts who must determine the location of a satellite during the station acquisition phase of a mission. Also described is the integration of four knowledge sources required to support the orbit determination assistant: orbital mechanics, spacecraft specifications, characteristics of the mission support software, and orbit analyst experience. This initial effort is continuing with expansion of YODA's capabilities, including evaluation of results of the orbit determination task.
Analysis of HY2A precise orbit determination using DORIS
NASA Astrophysics Data System (ADS)
Gao, Fan; Peng, Bibo; Zhang, Yu; Evariste, Ngatchou Heutchi; Liu, Jihua; Wang, Xiaohui; Zhong, Min; Lin, Mingsen; Wang, Nazi; Chen, Runjing; Xu, Houze
2015-03-01
HY2A is the first Chinese marine dynamic environment satellite. The payloads include a radar altimeter to measure the sea surface height in combination with a high precision orbit to be determined from tracking data. Onboard satellite tracking includes GPS, SLR, and the DORIS DGXX receiver which delivers phase and pseudo-range measurements. CNES releases raw phase and pseudo-range measurements with RINEX DORIS 3.0 format and pre-processed Doppler range-rate with DORIS 2.2 data format. However, the VMSI software package developed by Van Martin Systems, Inc which is used to estimate HY2A DORIS orbits can only process Doppler range-rate but not the DORIS phase data which are available with much shorter latency. We have proposed a method of constructing the phase increment data, which are similar to range-rate data, from RINEX DORIS 3.0 phase data. We compute the HY2A orbits from June, 2013 to August, 2013 using the POD strategy described in this paper based on DORIS 2.2 range-rate data and our reconstructed phase increment data. The estimated orbits are evaluated by comparing with the CNES precise orbits and SLR residuals. Our DORIS-only orbits agree with the precise GPS + SLR + DORIS CNES orbits radially at 1-cm and about 3-cm in the other two directions. SLR test with the 50° cutoff elevation shows that the CNES orbit can achieve about 1.1-cm accuracy in radial direction and our DORIS-only POD solutions are slightly worse. In addition, other HY2A DORIS POD concerns are discussed in this paper. Firstly, we discuss the frequency offset values provided with the RINEX data and find that orbit accuracy for the case when the frequency offset is applied is worse than when it is not applied. Secondly, HY2A DORIS antenna z-offsets are estimated using two kinds of measurements from June, 2013 to August, 2013. The results show that the measurement errors contribute a total of about 2-cm difference of estimated z-offset. Finally, we estimate HY2A orbits selecting 3 days with
Orbit Determination for Mars Global Surveyor During Mapping
NASA Technical Reports Server (NTRS)
Lemoine, F. G.; Rowlands, D. D.; Smith, D. E.; Pavlis, D. E.; Chinn, D. S.; Luthcke, S. B.; Neumann, G. A.
1999-01-01
The Mars Global Surveyor (MGS) spacecraft reached a low-altitude circular orbit on February 4, 1999, after the termination of the second phase of aerobraking. The MGS spacecraft carries the Mars Orbiter Laser Altimeter (MOLA) whose primary goal is to derive a global, geodetically referenced 0.2 deg x 0.2 deg topographic grid of Mars with a vertical accuracy of better than 30 meters. During the interim science orbits in the' Hiatus mission phase (October - November 1997), and the Science Phasing Orbits (March - April, 1998, and June - July 1998) 208 passes of altimeter data were collected by the MOLA instrument. On March 1, 1999 the first ten orbits of MOLA altimeter data from the near-circular orbit were successfully returned from MGS by the Deep Space Network (DSN). Data will be collected from MOLA throughout the Mapping phase of the MCS mission, or for at least one Mars year (687 days). Whereas the interim orbits of Hiatus and SPO were highly eccentric, and altimeter data were only collected near periapsis when the spacecraft was below 785 km, the Mapping orbit of MGS is near circular, and altimeter data will be collected continuously at a rate of 10 Hz. The proper analysis of the altimeter data requires that the orbit of the MGS spacecraft be known to an accuracy comparable to that of the quality of the altimeter data. The altimeter has an ultimate precision of 30 cm on mostly flat surfaces, so ideally the orbits of the MGS spacecraft should be known to this level. This is a stringent requirement, and more realistic goals of orbit error for MGS are ten to thirty meters. In this paper we will discuss the force and measurement modelling required to achieve this objective. Issues in force modelling include the proper modelling of the gravity field of Mars, and the modelling of non-conservatives forces, including the development of a 'macro-model', in a similar fashion to TOPEX/POSEIDON and TDRSS. During Cruise and Aerobraking, the high gain antenna (HGA) was stowed
Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter
NASA Technical Reports Server (NTRS)
Slojkowski, Steven; Lowe, Jonathan; Woodburn, James
2015-01-01
Orbit determination (OD) analysis results are presented for the Lunar Reconnaissance Orbiter (LRO) using a commercially available Extended Kalman Filter, Analytical Graphics' Orbit Determination Tool Kit (ODTK). Process noise models for lunar gravity and solar radiation pressure (SRP) are described and OD results employing the models are presented. Definitive accuracy using ODTK meets mission requirements and is better than that achieved using the operational LRO OD tool, the Goddard Trajectory Determination System (GTDS). Results demonstrate that a Vasicek stochastic model produces better estimates of the coefficient of solar radiation pressure than a Gauss-Markov model, and prediction accuracy using a Vasicek model meets mission requirements over the analysis span. Modeling the effect of antenna motion on range-rate tracking considerably improves residuals and filter-smoother consistency. Inclusion of off-axis SRP process noise and generalized process noise improves filter performance for both definitive and predicted accuracy. Definitive accuracy from the smoother is better than achieved using GTDS and is close to that achieved by precision OD methods used to generate definitive science orbits. Use of a multi-plate dynamic spacecraft area model with ODTK's force model plugin capability provides additional improvements in predicted accuracy.
An autonomous orbit determination method for MEO and LEO satellite
NASA Astrophysics Data System (ADS)
Zhang, Hui; Wang, Jin; Yu, Guobin; Zhong, Jie; Lin, Ling
2014-09-01
A reliable and secure navigation system and assured autonomous capability of satellite are in high demand in case of emergencies in space. This paper introduces a novel autonomous orbit determination method for Middle-Earth-Orbit and Low-Earth-Orbit (MEO and LEO) satellite by observing space objects whose orbits are known. Generally, the geodetic satellites, such as LAGEOS and ETALONS, can be selected as the space objects here. The precision CCD camera on tracking gimbal can make a series of photos of the objects and surrounding stars when MEO and LEO satellite encounters the space objects. Then the information processor processes images and attains sightings and angular observations of space objects. Several clusters of such angular observations are incorporated into a batch least squares filter to obtain an orbit determination solution. This paper describes basic principle and builds integrated mathematical model. The accuracy of this method is analyzed by means of computer simulation. Then a simulant experiment system is built, and the experimental results demonstrate the feasibility and effectiveness of this method. The experimental results show that this method can attain the accuracy of 150 meters with angular observations of 1 arcsecond system error.
49 CFR 7.31 - Initial determinations.
Code of Federal Regulations, 2010 CFR
2010-10-01
... 49 Transportation 1 2010-10-01 2010-10-01 false Initial determinations. 7.31 Section 7.31 Transportation Office of the Secretary of Transportation PUBLIC AVAILABILITY OF INFORMATION Time Limits § 7.31... to the life or physical safety of an individual; (ii) Requests made by a person primarily engaged...
50 CFR 296.9 - Initial determination.
Code of Federal Regulations, 2013 CFR
2013-10-01
... 50 Wildlife and Fisheries 11 2013-10-01 2013-10-01 false Initial determination. 296.9 Section 296.9 Wildlife and Fisheries NATIONAL MARINE FISHERIES SERVICE, NATIONAL OCEANIC AND ATMOSPHERIC ADMINISTRATION, DEPARTMENT OF COMMERCE CONTINENTAL SHELF FISHERMEN'S CONTINGENCY FUND § 296.9...
32 CFR 300.8 - Initial determinations.
Code of Federal Regulations, 2014 CFR
2014-07-01
... OF INFORMATION ACT PROGRAM DEFENSE LOGISTICS AGENCY FREEDOM OF INFORMATION ACT PROGRAM FOIA Request... withholding under one or more FOIA exemptions are met. DLA has IDAs throughout the agency; and each IDA will... 32 National Defense 2 2014-07-01 2014-07-01 false Initial determinations. 300.8 Section...
5 CFR 1631.7 - Initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 5 Administrative Personnel 3 2010-01-01 2010-01-01 false Initial determination. 1631.7 Section 1631.7 Administrative Personnel FEDERAL RETIREMENT THRIFT INVESTMENT BOARD AVAILABILITY OF RECORDS Production or Disclosure of Records Under the Freedom of Information Act, 5 U.S.C. 552 § 1631.7...
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 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.
Experimental determination of storage ring optics using orbit response measurements
NASA Astrophysics Data System (ADS)
Safranek, J.
1997-02-01
The measured response matrix giving the change in orbit at beam position monitors (BPMs) with changes in steering magnet excitation can be used to accurately calibrate the linear optics in an electron storage ring [1-8]. A computer code called LOCO (Linear Optics from Closed Orbits) was developed to analyze the NSLS X-Ray Ring measured response matrix to determine: the gradients in all 56 quadrupole magnets; the calibration of the steering magnets and BPMs; the roll of the quadrupoles, steering magnets, and BPMs about the electron beam direction; the longitudinal magnetic centers of the orbit steering magnets; the horizontal dispersion at the orbit steering magnets; and the transverse mis-alignment of the electron orbit in each of the sextupoles. Random orbit measurement error from the BPMs propagated to give only 0.04% rms error in the determination of individual quadrupole gradients and 0.4 mrad rms error in the determination of individual quadrupole rolls. Small variations of a few parts in a thousand in the quadrupole gradients within an individual family were resolved. The optics derived by LOCO gave accurate predictions of the horizontal dispersion, the beta functions, and the horizontal and vertical emittances, and it gave good qualitative agreement with the measured vertical dispersion. The improved understanding of the X-Ray Ring has enabled us to increase the synchrotron radiation brightness. The LOCO code can also be used to find the quadrupole family gradients that best correct for gradient errors in quadrupoles, in sextupoles, and from synchrotron radiation insertion devices. In this way the design periodicity of a storage ring's optics can be restored. An example of periodicity restoration will be presented for the NSLS VUV Ring. LOCO has also produced useful results when applied to the ALS storage ring [8].
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.
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
NASA Astrophysics Data System (ADS)
Yang, Yang; Yue, Xiaokui; Yuan, Jianping; Rizos, Chris
2014-11-01
Clock error estimation has been the focus of a great deal of research because of the extensive usage of clocks in GPS positioning applications. The receiver clock error in the spacecraft orbit determination is commonly estimated on an epoch-by-epoch basis, along with the spacecraft’s position. However, due to the high correlation between the spacecraft orbit altitude and the receiver clock parameters, estimates of the radial component are degraded in the kinematic approach. Using clocks with high stability, the predictable behaviour of the receiver oscillator can be exploited to improve the positioning accuracy, especially for the radial component. This paper introduces two GPS receiver clock models to describe the deterministic and stochastic property of the receiver clock, both of which can improve the accuracy of kinematic orbit determination for spacecraft in low earth orbit. In particular, the clock parameters are estimated as time offset and frequency offset in the two-state model. The frequency drift is also estimated as an unknown parameter in the three-state model. Additionally, residual non-deterministic random errors such as frequency white noise, frequency random walk noise and frequency random run noise are modelled. Test results indicate that the positioning accuracy could be improved significantly using one day of GRACE flight data. In particular, the error of the radial component was reduced by over 40.0% in the real-time scenario.
Automated Orbit Determination System (AODS) requirements definition and analysis
NASA Technical Reports Server (NTRS)
Waligora, S. R.; Goorevich, C. E.; Teles, J.; Pajerski, R. S.
1980-01-01
The requirements definition for the prototype version of the automated orbit determination system (AODS) is presented including the AODS requirements at all levels, the functional model as determined through the structured analysis performed during requirements definition, and the results of the requirements analysis. Also specified are the implementation strategy for AODS and the AODS-required external support software system (ADEPT), input and output message formats, and procedures for modifying the requirements.
Analysis on high-altitude earth Orbit Satellite Determination
NASA Astrophysics Data System (ADS)
He, J.; Hou, Y. W.; Yang, L.
2016-02-01
The difference is introduced between approx circular apogee orbit and approx circular perigee one by error transmitting at first. Then the characteristic of secant compensation is analysed when radar tracking object with high elevation. And two kinds of orbit force be pressed to, their perturbation influence and their earth-core angles are explained. And then the series of emulation results are shown including error data emulated with Monte Carlo method, the influence of the velocity increment from the ejecting force of spring while satellite-rocket separating and their perturbation influence and the length of influence of the data arc. Then decision analysis of Wald method and Bayesian statistics rule and the results from the two rule are introduced. So the suitable orbit determination decision is put forward from the decision method. Finally the result is tested reasonable and feasible via the real data. In the end it is useful to reference to make orbit decision in short injection of circular orbit far from the earth for calculating concurrently precise and timely.
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.
Hardware in-the-Loop Demonstration of Real-Time Orbit Determination in High Earth Orbits
NASA Technical Reports Server (NTRS)
Moreau, Michael; Naasz, Bo; Leitner, Jesse; Carpenter, J. Russell; Gaylor, Dave
2005-01-01
This paper presents results from a study conducted at Goddard Space Flight Center (GSFC) to assess the real-time orbit determination accuracy of GPS-based navigation in a number of different high Earth orbital regimes. Measurements collected from a GPS receiver (connected to a GPS radio frequency (RF) signal simulator) were processed in a navigation filter in real-time, and resulting errors in the estimated states were assessed. For the most challenging orbit simulated, a 12 hour Molniya orbit with an apogee of approximately 39,000 km, mean total position and velocity errors were approximately 7 meters and 3 mm/s respectively. The study also makes direct comparisons between the results from the above hardware in-the-loop tests and results obtained by processing GPS measurements generated from software simulations. Care was taken to use the same models and assumptions in the generation of both the real-time and software simulated measurements, in order that the real-time data could be used to help validate the assumptions and models used in the software simulations. The study makes use of the unique capabilities of the Formation Flying Test Bed at GSFC, which provides a capability to interface with different GPS receivers and to produce real-time, filtered orbit solutions even when less than four satellites are visible. The result is a powerful tool for assessing onboard navigation performance in a wide range of orbital regimes, and a test-bed for developing software and procedures for use in real spacecraft applications.
Orbit determination of satellite "Tance 1" with VLBI data
NASA Astrophysics Data System (ADS)
Huang, Yong; Hu, Xiao-gong; Huang, Cheng; Jiang, Dong-ro; Zheng, Wei-min; Zhang, Xiu-zhong
The satellite "Tance 1" of the "Double-Star Program" is the first truly scientific experimentation satellite of China. Its orbit is the farthest so far launched in China, with a geocentric apogee reaching 78 thousand kilometers. The tracking of "Tance 1" and of more distant space targets, such as the lunar exploration craft, can be realized with the VLBI technique of radio astronomy. In order to test and verify the role which the VLBI technique plays in the lunar exploration program of China, Shanghai Astronomical Observatory organized the only 3 tracking stations in China (located at Shanghai, Urumqi and Kunming), to carry out test tracking of "Tance 1," and used the time delay data obtained to determine the orbit of "Tance 1" over a two-day period, so providing a preliminary assessment of the possibility of VLBI orbit determination. The fitting error of the orbit so obtained is about 5.5 m in the time delay and about 2 cm/s in the delay rate (this for checking only), much better than is provided by the preliminary orbit (used merely for ensuring tracking) in which the corresponding figures are around 2 km and 15 cm/s. Further, if the orbit is determined by using both the time delay and time delay rate data (with weights according to their internal accuracies), then the residuals are 5.5 m in the time delay and 2 cm/s in the delay rate. For an appreciation of the true accuracy of the VLBI orbit determination, we used simulation data (of the observed two-day VLBI data) and found the results depended greatly on the error in the dynamic model of the satellite which, however, is difficult to assess, while the formal residuals are of the order of 1 kin in the delay and of cm/s in the delay rate. The simulation computation also indicates that a joint determination using both VLBI and USB data will have an improved accuracy.
Impact of Ionosphere on GPS-based Precise Orbit Determination of Low Earth Orbiters
NASA Astrophysics Data System (ADS)
Arnold, D.; Jaeggi, A.; Beutler, G.; Meyer, U.; Schaer, S.
2015-12-01
Deficiencies in geodetic products derived from the orbital trajectories of Low Earth Orbiting (LEO) satellites determined by GPS-based Precise Orbit Determination (POD) were identified in recent years. The precise orbits of the Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) mission are, e.g., severely affected by an increased position noise level over the geomagnetic poles and spurious signatures along the Earth's geomagnetic equator (see Fig. 1, which shows the carrier phase residuals of a reduced-dynamic orbit determination for GOCE in m). Such degradations may directly map into the gravity fields recovered from the orbits. They are related to a disturbed GPS signal propagation through the Earth's ionosphere and indicate that the GPS observation model and/or the data pre-processing need to be improved. While GOCE was the first mission where severe ionosphere-related problems became obvious, the GPS-based LEO POD of satellites of the more recent missions Swarm and Sentinel-1A turn out to be affected, as well. We characterize the stochastic and systematic behavior of the ionosphere by analyzing GPS data collected by the POD antennas of various LEO satellites covering a broad altitude range (e.g., GRACE, GOCE and Swarm) and for periods covering significant parts of an entire solar cycle, which probe substantially different ionosphere conditions. The information may provide the basis for improvements of data pre-processing to cope with the ionosphere-induced problems of LEO POD. The performance of cycle slip detection can, e.g., be degraded by large changes of ionospheric refraction from one measurement epoch to the next. Geographically resolved information on the stochastic properties of the ionosphere above the LEOs provide more realistic threshold values for cycle slip detection algorithms. Removing GPS data showing large ionospheric variations is a crude method to mitigate the ionosphere-induced artifacts in orbit and gravity field products
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.
Precise satellite orbit determination with particular application to ERS-1
NASA Astrophysics Data System (ADS)
Fernandes, Maria Joana Afonso Pereira
The motivation behind this study is twofold. First to assess the accuracy of ERS-1 long arc ephemerides using state of the art models. Second, to develop improved methods for determining precise ERS-1 orbits using either short or long arc techniques. The SATAN programs, for the computation of satellite orbits using laser data were used. Several facilities were added to the original programs: the processing of PRARE range and altimeter data, and a number of algorithms that allow more flexible solutions by adjusting a number of additional parameters. The first part of this study, before the launch of ERS-1, was done with SEAS AT data. The accuracy of SEASAT orbits computed with PRARE simulated data has been determined. The effect of temporal distribution of tracking data along the arc and the extent to which altimetry can replace range data have been investigated. The second part starts with the computation of ERS-1 long arc solutions using laser data. Some aspects of modelling the two main forces affecting ERS-l's orbit are investigated. With regard to the gravitational forces, the adjustment of a set of geopotential coefficients has been considered. With respect to atmospheric drag, extensive research has been carried out on determining the influence on orbit accuracy of the measurements of solar fluxes (P10.7 indices) and geomagnetic activity (Kp indices) used by the atmospheric model in the computation of atmospheric density at satellite height. Two new short arc methods have been developed: the Constrained and the Bayesian method. Both methods are dynamic and consist of solving for the 6 osculating elements. Using different techniques, both methods overcome the problem of normal matrix ill- conditioning by constraining the solution. The accuracy and applicability of these methods are discussed and compared with the traditional non-dynamic TAR method.
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
A reduced-dynamic technique for precise orbit determination
NASA Technical Reports Server (NTRS)
Wu, S. C.; Yunck, T. P.; Thornton, C. L.
1990-01-01
Observations of the Global Positioning System (GPS) will enable a reduced-dynamic technique for achieving subdecimeter orbit determination of earth-orbiting satellites. With this technique, information on the transition between satellite states at different observing times is furnished by both a formal dynamic model and observed satellite positional change (which is inferred kinematically from continuous GPS carrier-phase data). The relative weighting of dynamic and kinematic information can be freely varied. Covariance studies show that in situations where observing geometry is poor and the dynamic model is good, the model dominates determination of the state transition; where the dynamic model is poor and the geometry strong, carrier phase governs the determination of the transition. When neither kinematic nor dynamic information is clearly superior, the reduced-dynamic combination of the two can substantially improve the orbit-determination solution. Guidelines are given here for selecting a near-optimal weighting for the reduced-dynamic solution, and sensitivity of solution accuracy to this weighting is examined.
Orbital metastasis as initial manifestation of a widespread papillary thyroid microcarcinoma.
Pagsisihan, Daveric Ablis; Aguilar, Anthony Harvey Isabelo; Maningat, Ma Patricia Deanna Delfin
2015-01-01
Papillary thyroid carcinomas (PTCs), particularly microcarcinomas, rarely metastasise to the orbit. We report a case of a 49-year-old woman with a right supraorbital mass and unremarkable physical examination of the thyroid gland region. Orbital CT scan showed an expansile lytic lesion in the orbital plate of the frontal bone with a soft tissue component. An incision biopsy revealed metastatic well-differentiated thyroid carcinoma. Thyroid ultrasound was normal except for a subcentimetre nodule in the right lobe. The patient underwent total thyroidectomy where histopathology showed a subcentimetre follicular variant PTC. She subsequently received radioactive iodine therapy. Post-therapy whole body scan revealed metastatic thyroid tissues in the right orbital and posterior parietal, and left shoulder and hip areas. Although infrequent, metastatic thyroid carcinoma should be considered in patients with orbital metastasis even when neck examination is normal. In rare cases, this may be the initial manifestation of a widely metastatic papillary thyroid microcarcinoma. PMID:25819821
NASA Technical Reports Server (NTRS)
Taff, L. G.; Randall, P. M. S.
1985-01-01
A robust analytical formulation is developed to apply classical initial orbital determination to artificial satellites whose locations are uncertain to about 1 cu km and separated in time by no more than 30 min. An analytical simplification reduces Gauss's method, iteration on the semilatus rectum, iteration on the true anomaly, and the Lambert-Euler technique, to the solution of a single equation in one unknown, instead of the usual coupled triplet of three equations in three unknowns. The method is demonstrated for all common artificial satellite orbits over a variety of time intervals between the two location vectors, and for a varied set of position and distance errors.
Operational Challenges In TDRS Post-Maneuver Orbit Determination
NASA Technical Reports Server (NTRS)
Laing, Jason; Myers, Jessica; Ward, Douglas; Lamb, Rivers
2015-01-01
The GSFC Flight Dynamics Facility (FDF) is responsible for daily and post maneuver orbit determination for the Tracking and Data Relay Satellite System (TDRSS). The most stringent requirement for this orbit determination is 75 meters total position accuracy (3-sigma) predicted over one day for Terra's onboard navigation system. To maintain an accurate solution onboard Terra, a solution is generated and provided by the FDF Four hours after a TDRS maneuver. A number of factors present challenges to this support, such as maneuver prediction uncertainty and potentially unreliable tracking from User satellities. Reliable support is provided by comparing an extended Kalman Filter (estimated using ODTK) against a Batch Least Squares system (estimated using GTDS).
Meteoroid and Orbital Debris Threats to NASA's Docking Seals: Initial Assessment and Methodology
NASA Technical Reports Server (NTRS)
deGroh, Henry C., III; Gallo, Christopher A.; Nahra, Henry K.
2009-01-01
The Crew Exploration Vehicle (CEV) will be exposed to the Micrometeoroid Orbital Debris (MMOD) environment in Low Earth Orbit (LEO) during missions to the International Space Station (ISS) and to the micrometeoroid environment during lunar missions. The CEV will be equipped with a docking system which enables it to connect to ISS and the lunar module known as Altair; this docking system includes a hatch that opens so crew and supplies can pass between the spacecrafts. This docking system is known as the Low Impact Docking System (LIDS) and uses a silicone rubber seal to seal in cabin air. The rubber seal on LIDS presses against a metal flange on ISS (or Altair). All of these mating surfaces are exposed to the space environment prior to docking. The effects of atomic oxygen, ultraviolet and ionizing radiation, and MMOD have been estimated using ground based facilities. This work presents an initial methodology to predict meteoroid and orbital debris threats to candidate docking seals being considered for LIDS. The methodology integrates the results of ground based hypervelocity impacts on silicone rubber seals and aluminum sheets, risk assessments of the MMOD environment for a variety of mission scenarios, and candidate failure criteria. The experimental effort that addressed the effects of projectile incidence angle, speed, mass, and density, relations between projectile size and resulting crater size, and relations between crater size and the leak rate of candidate seals has culminated in a definition of the seal/flange failure criteria. The risk assessment performed with the BUMPER code used the failure criteria to determine the probability of failure of the seal/flange system and compared the risk to the allotted risk dictated by NASA s program requirements.
Meteoroid and Orbital Debris Threats to NASA's Docking Seals: Initial Assessment and Methodology
NASA Technical Reports Server (NTRS)
deGroh, Henry C., III; Nahra, Henry K.
2009-01-01
The Crew Exploration Vehicle (CEV) will be exposed to the Micrometeoroid Orbital Debris (MMOD) environment in Low Earth Orbit (LEO) during missions to the International Space Station (ISS) and to the micrometeoroid environment during lunar missions. The CEV will be equipped with a docking system which enables it to connect to ISS and the lunar module known as Altair; this docking system includes a hatch that opens so crew and supplies can pass between the spacecrafts. This docking system is known as the Low Impact Docking System (LIDS) and uses a silicone rubber seal to seal in cabin air. The rubber seal on LIDS presses against a metal flange on ISS (or Altair). All of these mating surfaces are exposed to the space environment prior to docking. The effects of atomic oxygen, ultraviolet and ionizing radiation, and MMOD have been estimated using ground based facilities. This work presents an initial methodology to predict meteoroid and orbital debris threats to candidate docking seals being considered for LIDS. The methodology integrates the results of ground based hypervelocity impacts on silicone rubber seals and aluminum sheets, risk assessments of the MMOD environment for a variety of mission scenarios, and candidate failure criteria. The experimental effort that addressed the effects of projectile incidence angle, speed, mass, and density, relations between projectile size and resulting crater size, and relations between crater size and the leak rate of candidate seals has culminated in a definition of the seal/flange failure criteria. The risk assessment performed with the BUMPER code used the failure criteria to determine the probability of failure of the seal/flange system and compared the risk to the allotted risk dictated by NASA's program requirements.
Orbit Determination Support for the Microwave Anisotropy Probe (MAP)
NASA Technical Reports Server (NTRS)
Truong, Son H.; Cuevas, Osvaldo O.; Slojkowski, Steven; Bauer, Frank H. (Technical Monitor)
2002-01-01
The Microwave Anisotropy Probe (MAP) is a Medium Class Explorers (MIDEX) mission produced in partnership between Goddard Space Flight Center (GSFC) and Princeton University. The main science objective of the MAP mission is to produce an accurate full-sky map of the cosmic microwave background temperature fluctuations anisotropy. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 2001. The spacecraft received a nominal direct insertion by the Delta into a 185 km circular orbit. MAP was then maneuvered into a sequence of phasing loops designed to set up a lunar swingby (gravity-assisted acceleration) of the spacecraft onto a transfer trajectory to a Lissajous orbit about the Earth-Sun L2 point. The mission duration is approximately 27 months with 3 to 4 months of transfer time to the final mission orbit about L2. The MAP transfer orbit consisted of 3.5 phasing loops: the first loop has a 7-day period, the second and third loops have a 9-day period, and the last half loop has a 4-day period as illustrated in Figure 1, which also indicates the placement of maneuvers. A Pfinal correction maneuver was performed 18 hours after the last perigee to more closely achieve the targeted lissajous orbit. The lunar encounter or swingby took place approximately 30 days after launch. After the lunar encounter, the spacecraft will cruise for approximately 120 days before it arrives at L2. A Mid-Course Correction (MCC) maneuver was executed seven days after the swingby to further refine the trajectory. Once the MAP satellite is injected into the L2 Lissajous orbit, it will perform occasional stationkeeping maneuvers to maintain the Lissajous orbit for a minimum of two years (and a goal of four years). Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes. Extensive trajectory error
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.
Determination of orbital drag perturbations caused by atmospheric effects
NASA Astrophysics Data System (ADS)
Sehnal, L.
Atmospheric perturbations of the elements of the artificial satellites orbits are determined using a special model of distribution and variations of the total density of the upper atmosphere between 200-500 km. The model includes diurnal and semi-annual density variations, variations with solar activity and geomagnetic index and latitudinal changes. The height profile is expressed by multiple exponential functions. The equations of motion are solved analytically.
Earth Observing System (EOS) real-time onboard orbit determination
NASA Technical Reports Server (NTRS)
Folta, David C.; Muller, Ron
1993-01-01
The paper describes the TDRSS Onboard Navigation System (TONS) selected by NASA/GSFC for the EOS-AM1 spacecraft as the baseline navigation system for real-time onboard orbit determination. Particular attention is given to the TONS algorithms and environmental models, the general design considerations, the algorithm implementation, and the required hardware. Results are presented of the covariance analysis for the nominal onboard and instrument requirements.
Gravity Recovery and Interior Laboratory Mission (GRAIL) Orbit Determination
NASA Technical Reports Server (NTRS)
You, Tung-Han; Antreasian, Peter; Broschart, Stephen; Criddle, Kevin; Higa, Earl; Jefferson, David; Lau, Eunice; Mohan, Swati; Ryne, Mark; Keck, Mason
2012-01-01
Launched on 10 September 2011 from the Cape Canaveral Air Force Station, Florida, the twin-spacecraft Gravity Recovery and Interior Laboratory (GRAIL) has the primary mission objective of generating a lunar gravity map with an unprecedented resolution via the Ka-band Lunar Gravity Ranging System (LGRS). After successfully executing nearly 30 maneuvers on their six-month journey, Ebb and Flow (aka GRAIL-A and GRAIL-B) established the most stringent planetary formation orbit on 1 March 2012 of approximately 30 km x 90 km in orbit size. This paper describes the orbit determination (OD) filter configurations, analyses, and results during the Trans-Lunar Cruise, Orbit Period Reduction, and Transition to Science Formation phases. The maneuver reconstruction strategies and their performance will also be discussed, as well as the navigation requirements, major dynamic models, and navigation challenges. GRAIL is the first mission to generate a full high-resolution gravity field of the only natural satellite of the Earth. It not only enables scientists to understand the detailed structure of the Moon but also further extends their knowledge of the evolutionary histories of the rocky inner planets. Robust and successful navigation was the key to making this a reality.
Galileo satellites measurement biases and orbit determination : preliminary results
NASA Astrophysics Data System (ADS)
Perosanz, Felix; Loyer, Sylvain; Mercier, Flavien; Boulanger, Cyrille; Capdeville, Hugues; Mezerette, Adrien
2013-04-01
Thanks to the IGS Multi-GNSS Experiment (M-GEX), signals from new GNSS satellites like Galileo are now available. CNES and IGN joined their efforts to contribute to the densification of this multi-GNSS global network through the REGINA project. However this network includes geodetic receivers from several manufacturers. For this reason we realized a dedicated test campaign to characterize the different receivers available in order to be able to process in a consistent way the data from the MGEX network. The test consisted in zero baseline measurements between receivers. Pseudo range as well as phase and wide-lane biases have been identified between Trimble, Leica, Javad and Septentrio receivers. Then the data from the global M-GEX tracking network have been processed for the Precise Orbit determination (POD) of the Galileo satellite. The strategy followed the one that the CNES-CLS IGS Analysis Center uses to compute hybrid GPS-GLONASS products. Since July 2012, Galileo data are processed and orbit solutions are routinely produced and evaluated. Pseudo-range and phase biases between receiver as well as inter-system biases have been quantified. We also demonstrated that a sub-decimeter 3D-WRMS orbit accuracy of Galileo satellite orbit can be achieved even during the constellation deployment.
Galileo satellites measurement biases and orbit determination: first results
NASA Astrophysics Data System (ADS)
Perosanz, F.; Loyer, S.; Mercier, F.; Boulanger, C.; Capdeville, H.; Mezerette, A.
2012-12-01
Thanks to the IGS Multi-GNSS Experiment (M-GEX), signals from new GNSS satellites like Galileo are now available. CNES and IGN joined their efforts to contribute to the densification of this multi-GNSS global network through the REGINA project. However this network includes geodetic receivers from several manufacturers. For this reason we realized a dedicated test campaign to characterize the different receivers available in order to be able to process in a consistent way the data from the MGEX network. The test consisted in zero baseline measurements between receivers. Pseudo range as well as phase and wide-lane biases have been identified between Trimble, Leica, Javad and Septentrio receivers. Then the data from the global M-GEX tracking network have been processed for the Precise Orbit determination of the Galileo satellite. The strategy followed the one that the CNES-CLS IGS Analysis Center uses to compute hybrid GPS-GLONASS products. Since July 2012, Galileo data are processed and orbit solutions are routinely produced and evaluated. Pseudo-range and phase biases between receiver as well as inter-system biases have been quantified. We also demonstrated that a decimeter 3D-WRMS orbit accuracy of Galileo satellite orbit can be achieved even during the constellation deployment.
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.
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.
20 CFR 405.120 - Effect of an initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Effect of an initial determination. 405.120 Section 405.120 Employees' Benefits SOCIAL SECURITY ADMINISTRATION ADMINISTRATIVE REVIEW PROCESS FOR ADJUDICATING INITIAL DISABILITY CLAIMS Initial Determinations § 405.120 Effect of an initial determination. An initial determination is binding...
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.; Sośnica, 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
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.
(42355) Typhon Echidna: Scheduling observations for binary orbit determination
NASA Astrophysics Data System (ADS)
Grundy, W. M.; Noll, K. S.; Virtanen, J.; Muinonen, K.; Kern, S. D.; Stephens, D. C.; Stansberry, J. A.; Levison, H. F.; Spencer, J. R.
2008-09-01
We describe a strategy for scheduling astrometric observations to minimize the number required to determine the mutual orbits of binary transneptunian systems. The method is illustrated by application to Hubble Space Telescope observations of (42355) Typhon-Echidna, revealing that Typhon and Echidna orbit one another with a period of 18.971±0.006 days and a semimajor axis of 1628±29 km, implying a system mass of (9.49±0.52)×10 kg. The eccentricity of the orbit is 0.526±0.015. Combined with a radiometric size determined from Spitzer Space Telescope data and the assumption that Typhon and Echidna both have the same albedo, we estimate that their radii are 76-16+14 and 42-9+8 km, respectively. These numbers give an average bulk density of only 0.44-0.17+0.44 gcm, consistent with very low bulk densities recently reported for two other small transneptunian binaries.
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.
Orbit determination for the final stages of the GOCE mission
NASA Astrophysics Data System (ADS)
Visser, Pieter N. A. M.; Van Den IJssel, Jose
The European Space Agency (ESA) Gravity field and steady-state Ocean Circulation Explorer (GOCE) re-entered the Earth's atmosphere on 11 November 2013, ending more than four years of successful mission operations. On 21 October 2013, the GOCE ion engines stopped functioning thereby ending its Drag-Free orbital motion. From then on, the satellite started decaying at increasing speeds and the atmospheric drag was growing exponentially towards its demise. Data from the GOCE Global Positioning System (GPS) receiver and accelerometers were collected down to an altitude of 137 km. The accelerometers started to get saturated on 7 November 2011 at an altitude of around 190 km, but kept on partially working until the altitude of 137 km as well. The data set of GPS and accelerometer observations for the last weeks of GOCE form a unique data set for testing GPS-based orbit determination in a high drag environment. Moreover, GPS-based estimates of the predominantly atmospheric drag induced non-gravitational accelerations can be validated against the accelerometer observations for this environment as well. This presentation will highlight precise orbit determination results and comparisons between accelerometer observations and GPS-based estimates of non-gravitational accelerations for the final stages of the GOCE mission.
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.
ANODE: An analytic orbit determination system, volume 1
NASA Astrophysics Data System (ADS)
Sridharan, R.; Seniw, W. P.
1980-06-01
The computer system at the Millstone Hill radar was upgraded in August, 1977, with the acquisition of a Harris 7/220 system. The new computer is a virtual memory multitasking system capable of supporting up to 768 K bytes of user program simultaneously. A software system design was made for the radar system with the new computer. One of the components of the system is an on-line real time analytic orbit determination program. The purpose of the program is threefold: (1) it is intended to act as a real-time monitor on the tracking performance of the radar; (2) it is designed to function as a rapid orbit estimator available interactively to analyst.
NASA Astrophysics Data System (ADS)
Li, R.; Oberst, J.; McEwen, A. S.; Archinal, B. A.; Beyer, R. A.; Thomas, P. C.; Chen, Y.; Hwangbo, J.; Lawver, J. D.; Scholten, F.; Mattson, S. S.; Howington-Kraus, A. E.; Robinson, M. S.
2009-12-01
The Lunar Reconnaissance Orbiter (LRO), launched June 18, 2009, carries the Lunar Reconnaissance Orbiter Camera (LROC) as one of seven remote sensing instruments on board. The camera system is equipped with a Wide Angle Camera (WAC) and two Narrow Angle Cameras (NAC) for systematic lunar surface mapping and detailed site characterization for potential landing site selection and resource identification. The LROC WAC is a pushframe camera with five 14-line by 704-sample framelets for visible light bands and two 16-line by 512-sample (summed 4x to 4 by 128) UV bands. The WAC can also acquire monochrome images with a 14-line by 1024-sample format. At the nominal 50-km orbit the visible bands ground scale is 75-m/pixel and the UV 383-m/pixel. Overlapping WAC images from adjacent orbits can be used to map topography at a scale of a few hundred meters. The two panchromatic NAC cameras are pushbroom imaging sensors each with a Cassegrain telescope of a 700-mm focal length. The two NAC cameras are aligned with a small overlap in the cross-track direction so that they cover a 5-km swath with a combined field-of-view (FOV) of 5.6°. At an altitude of 50-km, the NAC can provide panchromatic images from its 5,000-pixel linear CCD at a ground scale of 0.5-m/pixel. Calibration of the cameras was performed by using precision collimator measurements to determine the camera principal points and radial lens distortion. The orientation of the two NAC cameras is estimated by a boresight calibration using double and triple overlapping NAC images of the lunar surface. The resulting calibration results are incorporated into a photogrammetric bundle adjustment (BA), which models the LROC camera imaging geometry, in order to refine the exterior orientation (EO) parameters initially retrieved from the SPICE kernels. Consequently, the improved EO parameters can significantly enhance the quality of topographic products derived from LROC NAC imagery. In addition, an analysis of the spacecraft
Asteroid Orbit Determination and Rotational Period Calculations with CCD Astronomy
NASA Astrophysics Data System (ADS)
Burtz, Daniel C.
1998-10-01
This paper presents data collected and analyzed relating to photometry and astrometry of asteroids. All observations were accomplished at the U.S. Air Force Academy Observatory. The photometry involves determining the rotational period of asteroid 583 Klotilde. Astrometry was performed on asteroid 1035 Amata and the calculated position was used to determine its orbital elements. Klotilde was selected for rotational period determination based on its relatively low magnitude, favorable viewing position, and no previous rotational period information. Two hundred six images of Klotilde were taken and analyzed over four viewing nights. A Photometrics (PM512) Charge Couple Device (CCD) camera attached to a 61-cm Cassegrain telescope was used for these observations. Using NOAO' s IRAF software, the magnitudes of Klotilde and several comparison stars were determined. Using an Excel spreadsheet, differential photometry was performed and the light curve was plotted. The four nights of data gave a 9.210 +/- 0.005 hour synodic period with an amplitude of 0.18 magnitudes. Thirty-two images of Amata were taken on six different viewing nights. The images were taken with an ST-8 CCD attached to a 41-cm Cassegrain telescope. The data was reduced with the Astrometrica software package, which calculated the right ascension (RA), declination (Dec), and magnitude of Amata using several comparison stars. The computed RA and Dec, along with the times of observation were then used to determine the orbital elements of the asteroid.
HOW TO DETERMINE AN EXOMOON'S SENSE OF ORBITAL MOTION
Heller, René; Albrecht, Simon E-mail: albrecht@phys.au.dk
2014-11-20
We present two methods to determine an exomoon's sense of orbital motion (SOM), one with respect to the planet's circumstellar orbit and one with respect to the planetary rotation. Our simulations show that the required measurements will be possible with the European Extremely Large Telescope (E-ELT). The first method relies on mutual planet-moon events during stellar transits. Eclipses with the moon passing behind (in front of) the planet will be late (early) with regard to the moon's mean orbital period due to the finite speed of light. This ''transit timing dichotomy'' (TTD) determines an exomoon's SOM with respect to the circumstellar motion. For the 10 largest moons in the solar system, TTDs range between 2 and 12 s. The E-ELT will enable such measurements for Earth-sized moons around nearby Sun-like stars. The second method measures distortions in the IR spectrum of the rotating giant planet when it is transited by its moon. This Rossiter-McLaughlin effect (RME) in the planetary spectrum reveals the angle between the planetary equator and the moon's circumplanetary orbital plane, and therefore unveils the moon's SOM with respect to the planet's rotation. A reasonably large moon transiting a directly imaged planet like β 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.
Precision GPS orbit determination strategies for an earth orbiter and geodetic tracking system
NASA Technical Reports Server (NTRS)
Lichten, Stephen M.; Bertiger, Willy I.; Border, James S.
1988-01-01
Data from two 1985 GPS field tests were processed and precise GPS orbits were determined. With a combined carrier phase and pseudorange, the 1314-km repeatability improves substantially to 5 parts in 10 to the 9th (0.6 cm) in the north and 2 parts in 10 to the 8th (2-3 cm) in the other components. To achieve these levels of repeatability and accuracy, it is necessary to fine-tune the GPS solar radiation coefficients and ground station zenith tropospheric delays.
Improving GLONASS Precise Orbit Determination through Data Connection
Liu, Yang; Ge, Maorong; Shi, Chuang; Lou, Yidong; Wickert, Jens; Schuh, Harald
2015-01-01
In order to improve the precision of GLONASS orbits, this paper presents a method to connect the data segments of a single station-satellite pair to increase the observation continuity and, consequently, the strength of the precise orbit determination (POD) solution. In this method, for each GLONASS station-satellite pair, the wide-lane ambiguities derived from the Melbourne–Wübbena combination are statistically tested and corrected for phase integer offsets and then the same is carried out for the narrow-lane ambiguities calculated from the POD solution. An experimental validation was carried out using one-month GNSS data of a global network with 175 IGS stations. The result shows that, on average, 27.1% of the GLONASS station-satellite pairs with multiple data segments could be connected to a single long observation arc and, thus, only one ambiguity parameter was estimated. Using the connected data, the GLONASS orbit overlapping RMS at the day boundaries could be reduced by 19.2% in ideal cases with an averaged reduction of about 6.3%. PMID:26633414
Orbit determination covariance analysis for the Deep Space Program Science Experiment mission
NASA Technical Reports Server (NTRS)
Beckman, M.; Yee, C.; Lee, T.; Hoppe, M.; Oza, D.
1993-01-01
To define an appropriate orbit support procedure for the DSPSE mission, detailed permission orbit determination covariance analyses have been performed for the translunar and trans-Geographos mission phases. Preliminary analyses were also performed for the lunar mapping mission phase. These analyses are designed to assess the tracking patterns and the amount of tracking data needed to obtain orbit solutions of required accuracy for each mission phase and before and after each major orbit perturbation, such as orbit maneuvers and flybys of the Earth and Moon. In addition to operational orbit determination procedures, these analyses identify major error sources, estimate their contribution to orbital errors, and address possible strategies to reduce orbit determination error. For the lunar orbit phase, several lunar gravity error modeling approaches have been investigated. The covariance analysis results presented in this paper will serve as a guide for providing orbit determination support for the DSPSE mission.
Innovative observing strategy and orbit determination for Low Earth Orbit space debris
NASA Astrophysics Data System (ADS)
Milani, A.; Farnocchia, D.; Dimare, L.; Rossi, A.; Bernardi, F.
2012-03-01
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 (LEOs) region. The purpose is to determine the performances of a network of advanced optical sensors, through the use of the newest correlation and orbit determination algorithms. This network is foreseen for implementation in a Space Situational Awareness system, such as the future European one. 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 collision avoidance, and to detect catastrophic fragmentation events. The obtained results depend upon specific assumptions on the sensor and on the software technologies.
Relative Attitude Determination of Earth Orbiting Formations Using GPS Receivers
NASA Technical Reports Server (NTRS)
Lightsey, E. Glenn
2004-01-01
Satellite formation missions require the precise determination of both the position and attitude of multiple vehicles to achieve the desired objectives. In order to support the mission requirements for these applications, it is necessary to develop techniques for representing and controlling the attitude of formations of vehicles. A generalized method for representing the attitude of a formation of vehicles has been developed. The representation may be applied to both absolute and relative formation attitude control problems. The technique is able to accommodate formations of arbitrarily large number of vehicles. To demonstrate the formation attitude problem, the method is applied to the attitude determination of a simple leader-follower along-track orbit formation. A multiplicative extended Kalman filter is employed to estimate vehicle attitude. In a simulation study using GPS receivers as the attitude sensors, the relative attitude between vehicles in the formation is determined 3 times more accurately than the absolute attitude.
Initial on-orbit radiometric calibration of the Suomi NPP VIIRS reflective solar bands
NASA Astrophysics Data System (ADS)
Lei, Ning; Wang, Zhipeng; Fulbright, Jon; Lee, Shihyan; McIntire, Jeff; Chiang, Kwofu; Xiong, Xiaoxiong
2012-09-01
The on-orbit radiometric response calibration of the VISible/Near InfraRed (VISNIR) and the Short-Wave InfraRed (SWIR) bands of the Visible/Infrared Imager/Radiometer Suite (VIIRS) aboard the Suomi National Polar-orbiting Partnership (NPP) satellite is carried out through a Solar Diffuser (SD). The transmittance of the SD screen and the SD's Bidirectional Reflectance Distribution Function (BRDF) are measured before launch and tabulated, allowing the VIIRS sensor aperture spectral radiance to be accurately determined. The radiometric response of a detector is described by a quadratic polynomial of the detector's digital number (dn). The coefficients were determined before launch. Once on orbit, the coefficients are assumed to change by a common factor: the F-factor. The radiance scattered from the SD allows the determination of the F-factor. In this Proceeding, we describe the methodology and the associated algorithms in the determination of the F-factors and discuss the results.
A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer
NASA Technical Reports Server (NTRS)
Ikenaga, Toshinori; Hashida, Yoshi; Unwin, Martin
2007-01-01
In the present day, orbit determination by Global Positioning System (GPS) is not unusual. Especially for low-cost small satellites, position determination by an on-board GPS receiver provides a cheap, reliable and precise method. However, the original purpose of GPS is for ground users, so the transmissions from all of the GPS satellites are directed toward the Earth s surface. Hence there are some restrictions for users above the GPS constellation to detect those signals. On the other hand, a desire for precise orbit determination for users in orbits higher than GPS constellation exists. For example, the next Japanese Very Long Baseline Interferometry (VLBI) mission "ASTRO-G" is trying to determine its orbit in an accuracy of a few centimeters at apogee. The use of GPS is essential for such ultra accurate orbit determination. This study aims to construct a method for precise orbit determination for such high orbit users, especially in High Elliptical Orbits (HEOs). There are several approaches for this objective. In this study, a hybrid method with GPS and an accelerometer is chosen. Basically, while the position cannot be determined by an on-board GPS receiver or other Range and Range Rate (RARR) method, all we can do to estimate the user satellite s position is to propagate the orbit along with the force model, which is not perfectly correct. However if it has an accelerometer (ACC), the coefficients of the air drag and the solar radiation pressure applied to the user satellite can be updated and then the propagation along with the "updated" force model can improve the fitting accuracy of the user satellite s orbit. In this study, it is assumed to use an accelerometer available in the present market. The effects by a bias error of an accelerometer will also be discussed in this paper.
First Attempt of Orbit Determination of SLR Satellites and Space Debris Using Genetic Algorithms
NASA Astrophysics Data System (ADS)
Deleflie, F.; Coulot, D.; Descosta, R.; Fernier, A.; Richard, P.
2013-08-01
We present an orbit determination method based on genetic algorithms. Contrary to usual estimation methods mainly based on least-squares methods, these algorithms do not require any a priori knowledge of the initial state vector to be estimated. These algorithms can be applied when a new satellite is launched or for uncatalogued objects that appear in images obtained from robotic telescopes such as the TAROT ones. We show in this paper preliminary results obtained from an SLR satellite, for which tracking data acquired by the ILRS network enable to build accurate orbital arcs at a few centimeter level, which can be used as a reference orbit ; in this case, the basic observations are made up of time series of ranges, obtained from various tracking stations. We show as well the results obtained from the observations acquired by the two TAROT telescopes on the Telecom-2D satellite operated by CNES ; in that case, the observations are made up of time series of azimuths and elevations, seen from the two TAROT telescopes. The method is carried out in several steps: (i) an analytical propagation of the equations of motion, (ii) an estimation kernel based on genetic algorithms, which follows the usual steps of such approaches: initialization and evolution of a selected population, so as to determine the best parameters. Each parameter to be estimated, namely each initial keplerian element, has to be searched among an interval that is preliminary chosen. The algorithm is supposed to converge towards an optimum over a reasonable computational time.
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
Precise Orbit Determination of the GOCE Re-Entry Phase
NASA Astrophysics Data System (ADS)
Gini, Francesco; Otten, Michiel; Springer, Tim; Enderle, Werner; Lemmens, Stijn; Flohrer, Tim
2015-03-01
During the last days of the GOCE mission, after the GOCE spacecraft ran out of fuel, it slowly decayed before finally re-entering the atmosphere on the 11th November 2013. As an integrated part of the AOCS, GOCE carried a GPS receiver that was in operations during the re-entry phase. This feature provided a unique opportunity for Precise Orbit Determination (POD) analysis. As part of the activities carried out by the Navigation Support Office (HSO-GN) at ESOC, precise ephemerides of the GOCE satellite have been reconstructed for the entire re-entry phase based on the available GPS observations of the onboard LAGRANGE receiver. All the data available from the moment the thruster was switched off on the 21st of October 2013 to the last available telemetry downlink on the 10th November 2013 have been processed, for a total of 21 daily arcs. For this period a dedicated processing sequence has been defined and implemented within the ESA/ESOC NAvigation Package for Earth Observation Satellites (NAPEOS) software. The computed results show a post-fit RMS of the GPS undifferenced carrier phase residuals (ionospheric-free linear combination) between 6 and 14 mm for the first 16 days which then progressively increases up to about 80 mm for the last available days. An orbit comparison with the Precise Science Orbits (PSO) generated at the Astronomical Institute of the University of Bern (AIUB, Bern, Switzerland) shows an average difference around 9 cm for the first 8 daily arcs and progressively increasing up to 17 cm for the following days. During this reentry phase (21st of October - 10th November 2013) a substantial drop in the GOCE altitude is observed, starting from about 230 km to 130 km where the last GPS measurements were taken. During this orbital decay an increment of a factor of 100 in the aerodynamic acceleration profile is observed. In order to limit the mis-modelling of the non-gravitational forces (radiation pressure and aerodynamic effects) the newly developed
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.
Initial Mars Orbiter Laser Altimeter (MOLA) Measurements of the Mars Surface and Atmosphere
NASA Astrophysics Data System (ADS)
Abshire, James B.; Sun, Xiaoli; Afzal, Robert S.
1998-07-01
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.
Orbital metastasis as the initial presentation of lung adenocarcinoma: a case report
Sun, Liangchao; Qi, Yali; Sun, Xindong; Yu, Jinming; Meng, Xue
2016-01-01
Orbital metastasis as the initial presentation of lung adenocarcinoma is very rare, and so the lack of knowledge about this phenomenon can easily result in misdiagnosis, either as a orbital primary tumor or benign lesion. Here, we report a rare case in which the orbital symptom appeared first without any pulmonary manifestations. Our patient developed decreasing vision in his right eye over a 3-month duration. He then presented with proptosis and multiple aches from head to back. After systemic evaluation, our patient was diagnosed with Stage IV non-small-cell lung cancer and was managed with palliative chemoradiotherapy. The final result of treatment suggests that the therapeutic efficacy of chemotherapy on orbital metastasis is uncertain, and only some orbital metastatic masses may have a favorable response to radiation. Furthermore, we review the recent data and provide an in-depth discussion on the clinical features and course of ocular pulmonary metastases, and explain a new type of non-small-cell lung cancer metastatic pattern for ophthalmologists and oncologists to help them distinguish the orbital metastasis as the first manifestation. PMID:27274270
JASON-1 Precise Orbit Determination (POD)with SLR and DORIS Tracking
NASA Technical Reports Server (NTRS)
Zelensky, N. P.; Luthcke, S. B.; Rowlands, D. D.; Beckley, B. D.; Lemoine, Frank G.; Wang, Y. M.; Chinn, D. S.; Williams, T. A.
2002-01-01
Jason-1, the TOPEX/POSEIDON (T/P) radar altimeter follow-on, is intended to continue measurement of the ocean surface with the same, if not better accuracy. T/P has demonstrated that, the time variation of ocean topography can be determined with an accuracy of a few centimeters, thanks to the availability of highly accurate orbits based on SLR and DORIS tracking. For verification and cross-calibration, Jason-1, was initially injected into the T/P orbit, flying just 72 seconds ahead of T/P. This configuration lasted over 21 Jason cycles. In mid-August T/P was maneuvered into its final tandem configuration, a parallel groundtrack, in order to improve the combined coverage. Preliminary investigations using cycles 1-9, shown at the June 2002 SWT, indicated that nominal Jason orbits can achieve the 2-3 cm accuracy objective, however several puzzling aspects of SLR and DORIS measurement modeling were also observed. This paper presents recent analysis of Jason SLR+DORIS POD spanning more than 20 cycles, and revisits several of the more puzzling issues, including estimation of the Laser Retroreflector Array (LRA) offset. The accuracy of the orbits and of the measurement modeling are evaluated using several tests, including SLR, DORIS, and altimeter crossover residual analysis, altimeter collinear analysis, and direct comparison with GPS and other orbits. T/P POD results over the same period are used as a reference.
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)
Kibler, J. F.; Green, R. N.; Young, G. R.; Kelly, M. G.
1974-01-01
A method has previously been developed to satisfy terminal rendezvous and intermediate timing constraints for planetary missions involving orbital operations. The method uses impulse factoring in which a two-impulse transfer is divided into three or four impulses which add one or two intermediate orbits. The periods of the intermediate orbits and the number of revolutions in each orbit are varied to satisfy timing constraints. Techniques are developed to retarget the orbital transfer in the presence of orbit-determination and maneuver-execution errors. Sample results indicate that the nominal transfer can be retargeted with little change in either the magnitude (Delta V) or location of the individual impulses. Additonally, the total Delta V required for the retargeted transfer is little different from that required for the nominal transfer. A digital computer program developed to implement the techniques is described.
20 CFR 405.115 - Notice of the initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Notice of the initial determination. 405.115 Section 405.115 Employees' Benefits SOCIAL SECURITY ADMINISTRATION ADMINISTRATIVE REVIEW PROCESS FOR ADJUDICATING INITIAL DISABILITY CLAIMS Initial Determinations § 405.115 Notice of the initial determination. We will mail a written notice of...
NASA Technical Reports Server (NTRS)
Frauenholz, R. B.; Bhat, R. S.; Shapiro, B. E.; Leavitt, R. K.
1998-01-01
Since its' launch on August 10, 1992, the TOPEX/Poseidon satellite hs successfully observed the earth's ocean circulation using a combination of precision orbit determination (POD) and dual-frequency radar altimetry.
Urban Temperature Bias as Determined by Polar Orbiting Satellite Data
NASA Astrophysics Data System (ADS)
Johnson, Gregory Lynn
A method of determining urban temperature bias from remotely sensed data is developed and successfully tested in this study. First, atmospheric sounding products from NOAA's polar orbiting satellites were used to derive predictive equations of shelter-level maximum and minimum temperatures. Sounding data from both winter (January) and summer (July) months were combined with surface data from over 5300 cooperative weather stations in the continental United States to develop multiple linear regression equations. Predictive equations were then used to estimate rural ("background") temperatures, unaffected by urbanization. Clear and partly cloudy sounding retrievals proved superior to cloudy retrievals. Validation tests showed the models' abilities to predict rural temperatures in different months and in specific climatic regions. Using these equations, estimates of urban temperature bias for 37 cities in the United States were made. These estimates compared favorably to ground truth data. Largest differences between observed and predicted bias were found at coastal cities, and those at higher elevations in the western United States. Mean differences between observed and predicted bias for groups of cities were not significantly different, making the potential application of this technique to corrections of urban bias in large datasets very plausible. Other products obtained from polar orbiting satellites, including normalized difference vegetation index (NDVI) values, were also found to be useful descriptors of urban temperature bias. NDVI urban minus rural values were highly correlated to daily and monthly minimum temperature bias at most of the cities studied.
Improving integer ambiguity resolution for GLONASS precise orbit determination
NASA Astrophysics Data System (ADS)
Liu, Yang; Ge, Maorong; Shi, Chuang; Lou, Yidong; Wickert, Jens; Schuh, Harald
2016-05-01
The frequency division multiple access adopted in present GLONASS introduces inter-frequency bias (IFB) at the receiver-end both in code and phase observables, which makes GLONASS ambiguity resolution rather difficult or even not available, especially for long baselines up to several thousand kilometers. This is one of the major reasons that GLONASS could hardly reach the orbit precision of GPS, both in terms of consistency among individual International GNSS Service (IGS) analysis centers and discontinuity at the overlapping day boundaries. Based on the fact that the GLONASS phase IFB is similar on L1 and L2 bands in unit of length and is a linear function of the frequency number, several approaches have been developed to estimate and calibrate the IFB for integer ambiguity resolution. However, they are only for short and medium baselines. In this study, a new ambiguity resolution approach is developed for GLONASS global networks. In the approach, the phase ambiguities in the ionosphere-free linear combination are directly transformed with a wavelength of about 5.3 cm, according to the special frequency relationship of GLONASS L1 and L2 signals. After such transformation, the phase IFB rate can be estimated and corrected precisely and then the corresponding double-differenced ambiguities can be directly fixed to integers even for baselines up to several thousand kilometers. To evaluate this approach, experimental validations using one-month data of a global network with 140 IGS stations was carried out for GLONASS precise orbit determination. The results show that the GLONASS double-difference ambiguity resolution for long baselines could be achieved with an average fixing-rate of 91.4 %. Applying the fixed ambiguities as constraints, the GLONASS orbit overlapping RMS at the day boundaries could be reduced by 37.2 % in ideal cases and with an averaged reduction of about 21.4 %, which is comparable with that by the GPS ambiguity resolution. The orbit improvement is
Improving integer ambiguity resolution for GLONASS precise orbit determination
NASA Astrophysics Data System (ADS)
Liu, Yang; Ge, Maorong; Shi, Chuang; Lou, Yidong; Wickert, Jens; Schuh, Harald
2016-08-01
The frequency division multiple access adopted in present GLONASS introduces inter-frequency bias (IFB) at the receiver-end both in code and phase observables, which makes GLONASS ambiguity resolution rather difficult or even not available, especially for long baselines up to several thousand kilometers. This is one of the major reasons that GLONASS could hardly reach the orbit precision of GPS, both in terms of consistency among individual International GNSS Service (IGS) analysis centers and discontinuity at the overlapping day boundaries. Based on the fact that the GLONASS phase IFB is similar on L1 and L2 bands in unit of length and is a linear function of the frequency number, several approaches have been developed to estimate and calibrate the IFB for integer ambiguity resolution. However, they are only for short and medium baselines. In this study, a new ambiguity resolution approach is developed for GLONASS global networks. In the approach, the phase ambiguities in the ionosphere-free linear combination are directly transformed with a wavelength of about 5.3 cm, according to the special frequency relationship of GLONASS L1 and L2 signals. After such transformation, the phase IFB rate can be estimated and corrected precisely and then the corresponding double-differenced ambiguities can be directly fixed to integers even for baselines up to several thousand kilometers. To evaluate this approach, experimental validations using one-month data of a global network with 140 IGS stations was carried out for GLONASS precise orbit determination. The results show that the GLONASS double-difference ambiguity resolution for long baselines could be achieved with an average fixing-rate of 91.4 %. Applying the fixed ambiguities as constraints, the GLONASS orbit overlapping RMS at the day boundaries could be reduced by 37.2 % in ideal cases and with an averaged reduction of about 21.4 %, which is comparable with that by the GPS ambiguity resolution. The orbit improvement is
Analysis of filter tuning techniques for sequential orbit determination
NASA Technical Reports Server (NTRS)
Lee, T.; Yee, C.; Oza, D.
1995-01-01
This paper examines filter tuning techniques for a sequential orbit determination (OD) covariance analysis. Recently, there has been a renewed interest in sequential OD, primarily due to the successful flight qualification of the Tracking and Data Relay Satellite System (TDRSS) Onboard Navigation System (TONS) using Doppler data extracted onboard the Extreme Ultraviolet Explorer (EUVE) spacecraft. TONS computes highly accurate orbit solutions onboard the spacecraft in realtime using a sequential filter. As the result of the successful TONS-EUVE flight qualification experiment, the Earth Observing System (EOS) AM-1 Project has selected TONS as the prime navigation system. In addition, sequential OD methods can be used successfully for ground OD. Whether data are processed onboard or on the ground, a sequential OD procedure is generally favored over a batch technique when a realtime automated OD system is desired. Recently, OD covariance analyses were performed for the TONS-EUVE and TONS-EOS missions using the sequential processing options of the Orbit Determination Error Analysis System (ODEAS). ODEAS is the primary covariance analysis system used by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). The results of these analyses revealed a high sensitivity of the OD solutions to the state process noise filter tuning parameters. The covariance analysis results show that the state estimate error contributions from measurement-related error sources, especially those due to the random noise and satellite-to-satellite ionospheric refraction correction errors, increase rapidly as the state process noise increases. These results prompted an in-depth investigation of the role of the filter tuning parameters in sequential OD covariance analysis. This paper analyzes how the spacecraft state estimate errors due to dynamic and measurement-related error sources are affected by the process noise level used. This information is then used to establish
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.
NASA Astrophysics Data System (ADS)
Iona, Glenn; Butler, James; Guenther, Bruce; Graziani, Larissa; Johnson, Eric; Kennedy, Brian; Kent, Craig; Lambeck, Robert; Waluschka, Eugene; Xiong, Xiaoxiong
2012-09-01
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.
NASA Astrophysics Data System (ADS)
Bauer, S.; Hussmann, H.; Oberst, J.; Dirkx, D.; Mao, D.; Neumann, G. A.; Mazarico, E.; Torrence, M. H.; McGarry, J. F.; Smith, D. E.; Zuber, M. T.
2016-09-01
We used one-way laser ranging data from International Laser Ranging Service (ILRS) ground stations to NASA's Lunar Reconnaissance Orbiter (LRO) for a demonstration of orbit determination. In the one-way setup, the state of LRO and the parameters of the spacecraft and all involved ground station clocks must be estimated simultaneously. This setup introduces many correlated parameters that are resolved by using a priori constraints. Moreover the observation data coverage and errors accumulating from the dynamical and the clock modeling limit the maximum arc length. The objective of this paper is to investigate the effect of the arc length, the dynamical and modeling accuracy and the observation data coverage on the accuracy of the results. We analyzed multiple arcs using lengths of 2 and 7 days during a one-week period in Science Mission phase 02 (SM02, November 2010) and compared the trajectories, the post-fit measurement residuals and the estimated clock parameters. We further incorporated simultaneous passes from multiple stations within the observation data to investigate the expected improvement in positioning. The estimated trajectories were compared to the nominal LRO trajectory and the clock parameters (offset, rate and aging) to the results found in the literature. Arcs estimated with one-way ranging data had differences of 5-30 m compared to the nominal LRO trajectory. While the estimated LRO clock rates agreed closely with the a priori constraints, the aging parameters absorbed clock modeling errors with increasing clock arc length. Because of high correlations between the different ground station clocks and due to limited clock modeling accuracy, their differences only agreed at the order of magnitude with the literature. We found that the incorporation of simultaneous passes requires improved modeling in particular to enable the expected improvement in positioning. We found that gaps in the observation data coverage over 12 h (≈6 successive LRO orbits
Cassini Orbit Determination Performance (July 2008 - December 2011)
NASA Technical Reports Server (NTRS)
Pelletier, Frederic J.; Antreasian, Peter; Ardalan, Shadan; Buffington, Brent; Criddle, Kevin; Ionasescu, Rodica; Jacobson, Robert; Jones, Jeremy; Nandi, Sumita; Nolet, Simon; Parcher, Daniel; Roth, Duane; Smith, Jonathon; Thompson, Paul
2012-01-01
This paper reports on the orbit determination performance for the Cassini spacecraft from July 2008 to December 2011. During this period, Cassini made 85 revolutions around Saturn and had 52 close satellite encounters. 35 of those were with the massive Titan, 13 with the small, yet interesting, Enceladus as well as 2 with Rhea and 2 with Dione. The period also includes 4 double encounters, where engineers had to plan the trajectory for two close satellite encounters within days of each other at once. Navigation performance is characterized by ephemeris errors relative to in-flight predictions. Most Titan encounters 3-dimensional results are within a 1.5 formal sigma, with a few exceptions, mostly attributable to larger maneuver execution errors. Results for almost all other satellite encounter reconstructions are less than 3 sigma from their predictions. The errors are attributable to satellite ephemerides errors and in some cases to maneuver execution errors.
Orbit Determination Covariance Analysis for the Europa Clipper Mission
NASA Technical Reports Server (NTRS)
Ionasescu, Rodica; Martin-Mur, Tomas; Valerino, Powtawche; Criddle, Kevin; Buffington, Brent; McElrath, Timothy
2014-01-01
A new Jovian satellite tour is proposed by NASA, which would include numerous flybys of the moon Europa, and would explore its potential habitability by characterizing the existence of any water within and beneath Europa's ice shell. This paper describes the results of a covariance study that was undertaken on a sample tour to assess the navigational challenges and capabilities of such a mission from an orbit determination (OD) point of view, and to help establish a delta V budget for the maneuvers needed to keep the spacecraft on the reference trajectory. Additional parametric variations from the baseline case were also investigated. The success of the Europa Clipper mission will depend on the science measurements that it will enable. Meeting the requirements of the instruments onboard the spacecraft is an integral part of this analysis.
Determination of spin-orbit coefficients in semiconductor quantum wells
NASA Astrophysics Data System (ADS)
Faniel, S.; Matsuura, T.; Mineshige, S.; Sekine, Y.; Koga, T.
2011-03-01
We report the determination of the intrinsic spin-orbit interaction (SOI) parameters for In0.53Ga0.47As/In0.52Al0.48As quantum wells (QWs) from the analysis of the weak antilocalization effect. We show that the Dresselhaus SOI is mostly negligible in this system and that the intrinsic parameter for the Rashba effect, aSO≡α/
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.
NASA Astrophysics Data System (ADS)
Setty, Srinivas J.; Cefola, Paul J.; Montenbruck, Oliver; Fiedler, Hauke
2016-05-01
Catalog maintenance for Space Situational Awareness (SSA) demands an accurate and computationally lean orbit propagation and orbit determination technique to cope with the ever increasing number of observed space objects. As an alternative to established numerical and analytical methods, we investigate the accuracy and computational load of the Draper Semi-analytical Satellite Theory (DSST). The standalone version of the DSST was enhanced with additional perturbation models to improve its recovery of short periodic motion. The accuracy of DSST is, for the first time, compared to a numerical propagator with fidelity force models for a comprehensive grid of low, medium, and high altitude orbits with varying eccentricity and different inclinations. Furthermore, the run-time of both propagators is compared as a function of propagation arc, output step size and gravity field order to assess its performance for a full range of relevant use cases. For use in orbit determination, a robust performance of DSST is demonstrated even in the case of sparse observations, which is most sensitive to mismodeled short periodic perturbations. Overall, DSST is shown to exhibit adequate accuracy at favorable computational speed for the full set of orbits that need to be considered in space surveillance. Along with the inherent benefits of a semi-analytical orbit representation, DSST provides an attractive alternative to the more common numerical orbit propagation techniques.
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.
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.
Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter
NASA Technical Reports Server (NTRS)
Slojkowski, Steven; Lowe, Jonathan; Woodburn, James
2015-01-01
Since launch, the FDF has performed daily OD for LRO using the Goddard Trajectory Determination System (GTDS). GTDS is a batch least-squares (BLS) estimator. The tracking data arc for OD is 36 hours. Current operational OD uses 200 x 200 lunar gravity, solid lunar tides, solar radiation pressure (SRP) using a spherical spacecraft area model, and point mass gravity for the Earth, Sun, and Jupiter. LRO tracking data consists of range and range-rate measurements from: Universal Space Network (USN) stations in Sweden, Germany, Australia, and Hawaii. A NASA antenna at White Sands, New Mexico (WS1S). NASA Deep Space Network (DSN) stations. DSN data was sparse and not included in this study. Tracking is predominantly (50) from WS1S. The OD accuracy requirements are: Definitive ephemeris accuracy of 500 meters total position root-mean-squared (RMS) and18 meters radial RMS. Predicted orbit accuracy less than 800 meters root sum squared (RSS) over an 84-hour prediction span.
Initial On-Orbit Radiometric Calibration of the Suomi NPP VIIRS Reflective Solar Bands
NASA Technical Reports Server (NTRS)
Lei, Ning; Wang, Zhipeng; Fulbright, Jon; Lee, Shihyan; McIntire, Jeff; Chiang, Vincent; Xiong, Jack
2012-01-01
The on-orbit radiometric response calibration of the VISible/Near InfraRed (VISNIR) and the Short-Wave InfraRed (SWIR) bands of the Visible/Infrared Imager/Radiometer Suite (VIIRS) aboard the Suomi National Polar-orbiting Partnership (NPP) satellite is carried out through a Solar Diffuser (SD). The transmittance of the SD screen and the SD's Bidirectional Reflectance Distribution Function (BRDF) are measured before launch and tabulated, allowing the VIIRS sensor aperture spectral radiance to be accurately determined. The radiometric response of a detector is described by a quadratic polynomial of the detector?s digital number (dn). The coefficients were determined before launch. Once on orbit, the coefficients are assumed to change by a common factor: the F-factor. The radiance scattered from the SD allows the determination of the F-factor. In this Proceeding, we describe the methodology and the associated algorithms in the determination of the F-factors and discuss the results.
Orbital Moment Determination in (MnxFe1-x)3O4 Nanoparticles
Pool, V. L.; Jolley, C.; Douglas, T.; Arenholz, E.; Idzerda, Y. U.
2010-10-22
Nanoparticles of (Mn{sub x}Fe{sub 1-x}){sub 3}O{sub 4} with a concentration ranging from x = 0 to 1 and a crystallite size of 14-15 nm were measured using X-ray absorption spectroscopy and X-ray magnetic circular dichroism to determine the ratio of the orbital moment to the spin moment for Mn and Fe. At low Mn concentrations, the Mn substitutes into the host Fe{sub 3}O{sub 4} spinel structure as Mn{sup 2+} in the tetrahedral A-site. The net Fe moment, as identified by the X-ray dichroism intensity, is found to increase at the lowest Mn concentrations then rapidly decrease until no dichroism is observed at 20% Mn. The average Fe orbit/spin moment ratio is determined to initially be negative and small for pure Fe{sub 3}O{sub 4} nanoparticles and quickly go to 0 by 5%-10% Mn addition. The average Mn moment is anti-aligned to the Fe moment with an orbit/spin moment ratio of 0.12 which gradually decreases with Mn concentration.
An analytic development of orbit determination for a distant, planetary orbiter
NASA Technical Reports Server (NTRS)
Russell, R. K.; Thurman, S. W.
1989-01-01
With the advent of the Mariner '71 Mission, NASA has been sending spacecraft to orbit various distant bodies within the solar system. At present, there is still no adequate theory describing the inherent state estimation accuracy, based on two-way, coherent range-rate data. It is the purpose of this article to lay the groundwork for a general elliptic theory, and in addition to provide an analytic solution for the special case of circular orbits. It is shown that circular orbits about distant planets may suffer singularities in over-all position error estimation. These singularities are due to orbit inclination, placement of the line-of-nodes, and insignificant cross-velocity at the start and end of retrograde motion when orbiting a superior planet. Even though these conclusions appear to yield poor state estimation, one should not be unduly alarmed inasmuch as the stated conditions for singularity are not maintained for extended periods during typical mission scenarios. However, mission analysts should be aware of these potential pitfalls and realize that spuriously large results for circular orbiters can be obtained and are not the result of incorrect assumptions or faulty software. The general elliptic problem appears so involved that analytic inversion at this time is just not feasible, and in any case the resulting expression for the position error would likely be so lengthy that any understanding would be lost in the maze.
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.
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 ɛ~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.
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2012 CFR
2012-07-01
.... National Defense Other Regulations Relating to National Defense CENTRAL INTELLIGENCE AGENCY CHALLENGES TO... with this section. (b) Within 10 business days of receipt of a challenge, the Coordinator shall record... written response to a challenge within 60 business days of receipt. (d) If the C/CMCG determines that...
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2014 CFR
2014-07-01
.... National Defense Other Regulations Relating to National Defense CENTRAL INTELLIGENCE AGENCY CHALLENGES TO... with this section. (b) Within 10 business days of receipt of a challenge, the Coordinator shall record... written response to a challenge within 60 business days of receipt. (d) If the C/CMCG determines that...
32 CFR 1907.24. - Initial determination.
Code of Federal Regulations, 2013 CFR
2013-07-01
.... National Defense Other Regulations Relating to National Defense CENTRAL INTELLIGENCE AGENCY CHALLENGES TO... with this section. (b) Within 10 business days of receipt of a challenge, the Coordinator shall record... written response to a challenge within 60 business days of receipt. (d) If the C/CMCG determines that...
An Empirical State Error Covariance Matrix Orbit Determination Example
NASA Technical Reports Server (NTRS)
Frisbee, Joseph H., Jr.
2015-01-01
State estimation techniques serve effectively to provide mean state estimates. However, the state error covariance matrices provided as part of these techniques suffer from some degree of lack of confidence in their ability to adequately describe the uncertainty in the estimated states. A specific problem with the traditional form of state error covariance matrices is that they represent only a mapping of the assumed observation error characteristics into the state space. Any errors that arise from other sources (environment modeling, precision, etc.) are not directly represented in a traditional, theoretical state error covariance matrix. First, consider that an actual observation contains only measurement error and that an estimated observation contains all other errors, known and unknown. Then it follows that a measurement residual (the difference between expected and observed measurements) contains all errors for that measurement. Therefore, a direct and appropriate inclusion of the actual measurement residuals in the state error covariance matrix of the estimate will result in an empirical state error covariance matrix. This empirical state error covariance matrix will fully include all of the errors in the state estimate. The empirical error covariance matrix is determined from a literal reinterpretation of the equations involved in the weighted least squares estimation algorithm. It is a formally correct, empirical state error covariance matrix obtained through use of the average form of the weighted measurement residual variance performance index rather than the usual total weighted residual form. Based on its formulation, this matrix will contain the total uncertainty in the state estimate, regardless as to the source of the uncertainty and whether the source is anticipated or not. It is expected that the empirical error covariance matrix will give a better, statistical representation of the state error in poorly modeled systems or when sensor performance
NASA Astrophysics Data System (ADS)
Maione, F.; De Pietri, R.; Feo, A.; Löffler, F.
2016-09-01
We present results from three-dimensional general relativistic simulations of binary neutron star coalescences and mergers using public codes. We considered equal mass models where the baryon mass of the two neutron stars is 1.4{M}ȯ , described by four different equations of state (EOS) for the cold nuclear matter (APR4, SLy, H4, and MS1; all parametrized as piecewise polytropes). We started the simulations from four different initial interbinary distances (40,44.3,50, and 60 km), including up to the last 16 orbits before merger. That allows us to show the effects on the gravitational wave (GW) phase evolution, radiated energy and angular momentum due to: the use of different EOS, the orbital eccentricity present in the initial data and the initial separation (in the simulation) between the two stars. Our results show that eccentricity has a major role in the discrepancy between numerical and analytical waveforms until the very last few orbits, where ‘tidal’ effects and missing high-order post-Newtonian coefficients also play a significant role. We test different methods for extrapolating the GW signal extracted at finite radii to null infinity. We show that an effective procedure for integrating the Newman–Penrose {\\psi }4 signal to obtain the GW strain h is to apply a simple high-pass digital filter to h after a time domain integration, where only the two physical motivated integration constants are introduced. That should be preferred to the more common procedures of introducing additional integration constants, integrating in the frequency domain or filtering {\\psi }4 before integration.
Procedure for the Determination of Orbits of Astronomical Bodies
ERIC Educational Resources Information Center
Birnbaum, David
1977-01-01
Presents a procedure for finding the elements of the orbit of an astronomical object from three or more observations. From a set of assumed elements an ephemeris is calculated and compared to the observations. (MLH)
In-Orbit Lifetime Prediction for LEO and HEO Based on Orbit Determination from TLE Data
NASA Astrophysics Data System (ADS)
Agueda, A.; Aivar, L.; Tirado, J.; Dolado, J. C.
2013-08-01
Objects in Low-Earth Orbits (LEO) and Highly Elliptical Orbits (HEO) are subjected to decay and re-entry into the atmosphere due mainly to the drag force. While being this process the best solution to avoid the proliferation of debris in space and ensure the sustainability of future space activities, it implies a threat to the population on ground. Thus, the prediction of the in-orbit lifetime of an object and the evaluation of the risk on population and ground assets constitutes a crucial task. This paper will concentrate on the first of these tasks. Unfortunately the lifetime of an object in space is remarkably difficult to predict. This is mainly due to the dependence of the atmospheric drag on a number of uncertain elements such as the density profile and its dependence on the solar activity, the atmospheric conditions, the mass and surface area of the object (very difficult to evaluate), its uncontrolled attitude, etc. In this paper we will present a method for the prediction of this lifetime based on publicly available Two-Line Elements (TLEs) from the American USSTRATCOM's Joint Space Operations Center (JSpOC). TLEs constitute an excellent source to access routinely orbital information for thousands of objects even though of their reduced and unpredictable accuracy. Additionally, the implementation of the method on a CNES's Java-based tool will be presented. This tool (OPERA) is executed routinely at CNES to predict the orbital lifetime of a whole catalogue of objects.
20 CFR 320.10 - Reconsideration of initial determination.
Code of Federal Regulations, 2013 CFR
2013-04-01
... 20 Employees' Benefits 1 2013-04-01 2012-04-01 true Reconsideration of initial determination. 320.10 Section 320.10 Employees' Benefits RAILROAD RETIREMENT BOARD REGULATIONS UNDER THE RAILROAD UNEMPLOYMENT INSURANCE ACT INITIAL DETERMINATIONS UNDER THE RAILROAD UNEMPLOYMENT INSURANCE ACT AND REVIEWS OF AND APPEALS FROM SUCH DETERMINATIONS...
An Independent Orbit Determination Simulation for the OSIRIS-REx Asteroid Sample Return Mission
NASA Technical Reports Server (NTRS)
Getzandanner, Kenneth; Rowlands, David; Mazarico, Erwan; Antreasian, Peter; Jackman, Coralie; Moreau, Michael
2016-01-01
After arriving at the near-Earth asteroid (101955) Bennu in late 2018, the OSIRIS-REx spacecraft will execute a series of observation campaigns and orbit phases to accurately characterize Bennu and ultimately collect a sample of pristine regolith from its surface. While in the vicinity of Bennu, the OSIRIS-REx navigation team will rely on a combination of ground-based radiometric tracking data and optical navigation (OpNav) images to generate and deliver precision orbit determination products. Long before arrival at Bennu, the navigation team is performing multiple orbit determination simulations and thread tests to verify navigation performance and ensure interfaces between multiple software suites function properly. In this paper, we will summarize the results of an independent orbit determination simulation of the Orbit B phase of the mission performed to test the interface between the OpNav image processing and orbit determination software packages.
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
Orbits of four visual binaries determined from observations along short arcs
NASA Astrophysics Data System (ADS)
Romanenko, L. G.; Kiselev, A. A.
2014-01-01
The orbits of the four visual binaries ADS 246 (GL 15), ADS 7724 ( γ Leo), ADS 10386 (GJ 659), and ADS 14909 (1 Peg) have been determined using the apparent motion parameters (AMP) method. The orbital periods of these stars are 1200, 550, 7500, and 18 000 yr, respectively. The orbits were calculated based on observations along short arcs obtained with the 26-inch refractor of the Pulkovo Observatory and Hipparcos parallaxes, supplemented with radial-velocity measurements for the components of these pairs taken from the literature. All visual and photographic observations of these stars after 1830 from the WDS catalog have been taken into consideration. The new orbits of ADS 246 and ADS 7724 are compared with the orbits computed in other studies. The orbits of ADS 10386 and ADS 14909 have been determined for the first time. The orientation of the planes of the computed orbits in Galactic coordinates have also been calculated.
Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise
NASA Technical Reports Server (NTRS)
Pachura, Daniel A.; Vavrina, Matthew A.; Carpenter, J. R.; Wright, Cinnamon A.
2014-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the missions commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.
Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise
NASA Technical Reports Server (NTRS)
Pachura, Daniel A.; Vavrina, Matthew A.; Carpenter, J. Russell; Wright, Cinnamon A.
2014-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the mission's commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.
Orbit determination and analysis of meteors recently observed by Finnish Fireball Network
NASA Astrophysics Data System (ADS)
Dmitriev, V.; Lupovla, V.; Gritsevich, M.; Lyytinen, E.; Mineeva, S.
2015-10-01
We perform orbit determination and analysis of three fireballs recently observed by Finnish Fireball Network (FFN). Precise orbit determination was performed by using integration of differential equations of motion. This technique was implemented into free distributable software "Meteor Toolkit". Accounting of several perturbing forces are discussed. Also estimation of accuracy of orbital elements was obtained by propagation of observational error with using covariance transformation. Long-term backward integration was provided as well.
Plasma Cell Neoplasm Manifesting Initially as a Sub-Cutaneous Supra-Orbital Swelling.
Jaiswal, Riddhi; Agarwal, Garima; Singh, Sudhir
2016-01-01
Multiple myeloma is a plasma cell neoplasm seen usually in patients over 50 years of age. Some cases may be asymptomatic initially and are detected during a routine test like complete blood count. They only require a close follow-up and monitoring. However, around 1% of these monoclonal gammopathy of undetermined significance progress to multiple myeloma every year and then they need to be taken care of by chemotherapy, targeted therapy, bisphosphonates and 6 monthly urine and bone examinations. Here, we present a case of 35-year-old female with an initial symptom of a vague backache along with a left subcutaneous supra-orbital swelling which was diagnosed as multiple myeloma by aspiration cytology and confirmed by ancillary tests. She has since been on treatment with bortezomib and prednisone and is responding well. PMID:26955130
Plasma Cell Neoplasm Manifesting Initially as a Sub-Cutaneous Supra-Orbital Swelling
Jaiswal, Riddhi; Agarwal, Garima; Singh, Sudhir
2016-01-01
Multiple myeloma is a plasma cell neoplasm seen usually in patients over 50 years of age. Some cases may be asymptomatic initially and are detected during a routine test like complete blood count. They only require a close follow-up and monitoring. However, around 1% of these monoclonal gammopathy of undetermined significance progress to multiple myeloma every year and then they need to be taken care of by chemotherapy, targeted therapy, bisphosphonates and 6 monthly urine and bone examinations. Here, we present a case of 35-year-old female with an initial symptom of a vague backache along with a left subcutaneous supra-orbital swelling which was diagnosed as multiple myeloma by aspiration cytology and confirmed by ancillary tests. She has since been on treatment with bortezomib and prednisone and is responding well. PMID:26955130
Impact of the ionosphere on GPS-based precise orbit determination of Low Earth Orbiters
NASA Astrophysics Data System (ADS)
Arnold, Daniel; Jäggi, Adrian; Meyer, Ulrich; Beutler, Gerhard
2016-04-01
GPS-derived kinematic precise Swarm orbits are significantly affected by increased position noise over the geomagnetic poles and spurious signatures along the geomagnetic equator. The latter deficiencies were identified for the first time for the Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) mission and are attributed to the distortion of the GPS carrier signal when propagating through portions of the Earth's ionosphere with a large free electron content. Via the GPS-derived kinematic Swarm positions, the spurious signatures along the geomagnetic equator map directly into the derived gravity fields. This was already the case for GOCE and obviously is also true for Swarm. To identify the root cause of the problem, the stochastic and deterministic behavior of the ionosphere is characterized by analyzing data collected by the GPS receivers on various LEO satellites. We compare in particular the performance of the Swarm and the GRACE receivers, because no obvious degradations occur in GRACE orbit and gravity field solutions. Removing GPS data with large ionospheric variations mitigates the ionosphere-induced artifacts in orbits and gravity fields. We illustrate the impact of this measure on the Swarm orbit and gravity field solutions. Making use of the geographically resolved ionosphere characteristics, e.g., to establish better data weighting schemes, results in a better POD performance for LEO satellites.
U Geminorum: a Test Case for Orbital Parameters Determination
NASA Astrophysics Data System (ADS)
Echevarría, Juan; de La Fuente, Eduardo; Costero, Rafael
2007-08-01
Due to its eclipsing nature and thorough observational studies, U Gem, in general, a good candidate for the analysis of standard and new methods in the determination of the orbital parameters in cataclysmic variables. Although in this interactive binary, these parameters are relatively well known, there are still discrepancies in the radial velocity semi-amplitude of the white dwarf, as obtained from the optical or the Ultraviolet data. Furthermore, the secondary star is not visible in the optical; consequently, its corresponding semi-amplitude has been derived from data obtained in the infrared region. For these reasons U Gem is an interesting case for testing new methods to derive orbital parameters based on optical observations only. High resolution spectroscopy of U Gem, covering the spectral region λ 5200-9100 Å, was obtained. The system was observed during quiescence, about 35 days after the onset of an outburst. We did not find a hot spot or gas stream around the outer boundaries of the accretion disk. Instead, we detected a strong narrow emission at the location of the secondary star. We measured the radial velocity curve from the wings of the double-peaked Hα emission line, and obtained a semi-amplitude value in excellent agreement with the ultraviolet results by Long & Gilliland (1999). We present also a new method to obtain K[2], based on the detection of the TiO band around λ 7050 Å. Our results are compared with published values derived from the near-infrared NaI line doublet. From a comparison of the TiO band with those of late type M stars, we find that a best fit is obtained for a M6 V star, contributing 5 percent of the total light at that spectral region. Assuming that the radial velocity semi-amplitudes reflect accurately the motion of the binary components, then from our results: K[em] = 108 km s-1 and K[abs] = 310 km s-1. For a revised inclination angle of i = 70o (Zhang et al. 1987) the system parameters become; M[wd] = 1.20 ± 0.05 M
Dawn Orbit Determination Team: Modeling and Fitting of Optical Data at Vesta
NASA Technical Reports Server (NTRS)
Kennedy, Brian; Abrahamson, Matt; Ardito, Alessandro; Haw, Robert; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The Dawn spacecraft was launched on September 27th, 2007. Its mission is to consecutively rendezvous with and observe the two largest bodies in the main asteroid belt, Vesta and Ceres. It has already completed over a year's worth of direct observations of Vesta (spanning from early 2011 through late 2012) and is currently on a cruise trajectory to Ceres, where it will begin scientific observations in mid-2015. Achieving this data collection required careful planning and execution from all Dawn operations teams. Dawn's Orbit Determination (OD) team was tasked with reconstruction of the as-flown trajectory as well as determination of the Vesta rotational rate, pole orientation and ephemeris, among other Vesta parameters. Improved knowledge of the Vesta pole orientation, specifically, was needed to target the final maneuvers that inserted Dawn into the first science orbit at Vesta. To solve for these parameters, the OD team used radiometric data from the Deep Space Network (DSN) along with optical data reduced from Dawn's Framing Camera (FC) images. This paper will de-scribe the initial determination of the Vesta ephemeris and pole using a combination of radiometric and optical data, and also the progress the OD team has made since then to further refine the knowledge of Vesta's body frame orientation and rate with these data.
Real-Time and Post-Processed Orbit Determination and Positioning
NASA Technical Reports Server (NTRS)
Bar-Sever, Yoaz E. (Inventor); Bertiger, William I. (Inventor); Dorsey, Angela R. (Inventor); Harvey, Nathaniel E. (Inventor); Lu, Wenwen (Inventor); Miller, Kevin J. (Inventor); Miller, Mark A. (Inventor); Romans, Larry J. (Inventor); Sibthorpe, Anthony J. (Inventor); Weiss, Jan P. (Inventor); Garcia Fernandez, Miquel (Inventor); Gross, Jason (Inventor)
2016-01-01
Novel methods and systems for the accurate and efficient processing of real-time and latent global navigation satellite systems (GNSS) data are described. Such methods and systems can perform orbit determination of GNSS satellites, orbit determination of satellites carrying GNSS receivers, positioning of GNSS receivers, and environmental monitoring with GNSS data.
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.
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.
The iterative solution of the problem of orbit determination using Chebyshev series
NASA Technical Reports Server (NTRS)
Feagin, T.
1975-01-01
A method of orbit determination is investigated which employs Picard iteration and Chebyshev series. The method is applied to the problem of determining the orbit of an earth satellite from range and range-rate observations contaminated by noise. It is shown to be readily applicable and to possess linear convergence.
Program of Aes Orbit Determination from Measurement Data of Astronomical Station ("orbita - M")
NASA Astrophysics Data System (ADS)
Sheptoon, A. D.; Kolesnik, S. Ja.; Paltsev, N. G.
A program is developed of determining AES orbits from measurement data of one or several astronomical stations. Its algorithm is rather stable to small errors of measurements and permits to use data with low accuracy for calculations.The use of several transits data enables to increase presision of orbital semi-major axe determination by nearly 10000 times.
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.
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.
Initial Data for Binary Neutron Stars with Arbitrary Spin and Orbital Eccentricity
NASA Astrophysics Data System (ADS)
Tsatsin, Petr; Marronetti, Pedro
2013-04-01
The starting point of any general relativistic numerical simulation is a solution of the Hamiltonian and momentum constraint. One characteristic of the Binary Neutron Star (BNS) initial data problem is that, unlike the case of binary black holes, there are no formalisms that permit the construction of initial data for stars with arbitrary spins. For many years, the only options available have been systems either with irrotational or corotating fluid. Ten years ago, Marronetti & Shapiro (2003) introduced an approximation that would produce such arbitrarily spinning systems. More recently, Tichy (2012) presented a new formulation to do the same. However, all these data sets are bound to have a non-zero eccentricity that results from the fact the stars' velocity have initial null radial components. We present here a new approximation for BNS initial data for systems that possess arbitrary spins and arbitrary radial and tangential velocity components. The latter allows for the construction of data sets with arbitrary orbital eccentricity. Through the fine-tuning of the radial component, we were able to reduce the eccentricity by a factor of several compared to that of standard helical symmetry data sets such as those currently used in the scientific community.
Validation of individual GOCE accelerometers by precise orbit determination
NASA Astrophysics Data System (ADS)
Visser, Pieter N. A. M.
2012-07-01
The European Space Agency (ESA) 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 (SST) observations by the Global Positioning System (GPS) and which are claimed to be at the few cm precision level, can be used to 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 coordinates 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 opposed to using the so-called common-mode accelerations, provided the center of the gradiometer coincides with the center of mass. Dynamic orbit fits based on these common-mode accelerations therefore served as reference. It will be shown that for all individual accelerometers similar dynamic orbit fits can be obtained, provided the above mentioned corrections are made. When using the common-mode accelerations, similar fits are obtained. In addition, attention will be paid to the possibility of estimating accelerometer calibration parameters, such as biases and scale factors.
Orbit determination of highly elliptical Earth orbiters using improved Doppler data-processing modes
NASA Technical Reports Server (NTRS)
Estefan, J. A.
1995-01-01
A navigation error covariance analysis of four highly elliptical Earth orbits is described, with apogee heights ranging from 20,000 to 76,800 km and perigee heights ranging from 1,000 to 5,000 km. This analysis differs from earlier studies in that improved navigation data-processing modes were used to reduce the radio metric data. For this study, X-band (8.4-GHz) Doppler data were assumed to be acquired from two Deep Space Network radio antennas and reconstructed orbit errors propagated over a single day. Doppler measurements were formulated as total-count phase measurements and compared to the traditional formulation of differenced-count frequency measurements. In addition, an enhanced data-filtering strategy was used, which treated the principal ground system calibration errors affecting the data as filter parameters. Results suggest that a 40- to 60-percent accuracy improvement may be achievable over traditional data-processing modes in reconstructed orbit errors, with a substantial reduction in reconstructed velocity errors at perigee. Historically, this has been a regime in which stringent navigation requirements have been difficult to meet by conventional methods.
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.
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2014 CFR
2014-04-01
... 18 Conservation of Power and Water Resources 2 2014-04-01 2014-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2013 CFR
2013-04-01
... 18 Conservation of Power and Water Resources 2 2013-04-01 2012-04-01 true Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2012 CFR
2012-04-01
... 18 Conservation of Power and Water Resources 2 2012-04-01 2012-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Initial determinations of SSI eligibility... INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of SSI eligibility. (a) What happens when you apply for SSI benefits. When you apply for SSI benefits we will ask...
18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
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18 CFR 701.309 - Appeal of initial adverse determination.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 18 Conservation of Power and Water Resources 2 2011-04-01 2011-04-01 false Appeal of initial adverse determination. 701.309 Section 701.309 Conservation of Power and Water Resources WATER RESOURCES COUNCIL COUNCIL ORGANIZATION Protection of Privacy § 701.309 Appeal of initial adverse determination....
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2014 CFR
2014-04-01
... 20 Employees' Benefits 2 2014-04-01 2014-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2011 CFR
2011-04-01
... 20 Employees' Benefits 2 2011-04-01 2011-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2013 CFR
2013-04-01
... 20 Employees' Benefits 2 2013-04-01 2013-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
20 CFR 416.203 - Initial determinations of SSI eligibility.
Code of Federal Regulations, 2012 CFR
2012-04-01
... 20 Employees' Benefits 2 2012-04-01 2012-04-01 false Initial determinations of SSI eligibility. 416.203 Section 416.203 Employees' Benefits SOCIAL SECURITY ADMINISTRATION SUPPLEMENTAL SECURITY INCOME FOR THE AGED, BLIND, AND DISABLED Eligibility General § 416.203 Initial determinations of...
Dawn Orbit Determination Team : Trajectory Modeling and Reconstruction Processes at Vesta
NASA Technical Reports Server (NTRS)
Abrahamson, Matt; Ardito, Alessandro; Han, Don; Haw, Robert; Kennedy, Brian; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The NASA Dawn spacecraft was launched on September 27, 2007 on a mission to study the asteroid belt's two largest objects, Vesta and Ceres. It is the first deep space orbiting mission to demonstrate solar-electric ion propulsion, providing the necessary delta-V to enable capture and escape from two extraterrestrial bodies. At this time, Dawn has completed its science campaign at Vesta and is currently on its journey to Ceres, where it will arrive in mid-2015. The spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012, capturing science data during four dedicated orbit phases. In order to maintain the reference orbits necessary for science and enable the transfers between those orbits, precise and timely orbit determination was required. The constraints associated with low-thrust ion propulsion coupled with the relatively unknown a priori gravity and rotation models for Vesta presented unique challenges for the Dawn orbit determination team. While [1] discusses the prediction performance of the orbit determination products, this paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.
Preliminary Orbit Determination of a Tethered Satellite Using the p-Iteration Method
NASA Astrophysics Data System (ADS)
Cicci, D. A.; Qualls, C.
2016-06-01
The possibility of the future deployment of tethered satellites has created a need for a preliminary orbit determination method capable of determining whether or not a satellite is tethered to another satellite. Classical preliminary orbit determination methods, which are used for untethered satellites, typically require two or more position vectors along with their respective observation times in order to determine a preliminary orbital element set. However, these conventional methods can't distinguish between an untethered satellite and a tethered one, whose motion is modified due to the presence of a tether force. The use of conventional methods on a satellite which is part of a tethered satellite system will result in the calculation of inaccurate orbital elements. Modifications have been made to the p-iteration preliminary orbit determination method in order to allow for the identification of these tethered satellites. The modifications allow for the calculation of a modified gravitational parameter, which can be used to distinguish between a tethered satellite and an untethered one. This paper applies this modified p-iteration method to the problem of the quick identification of a tethered satellite. The performance of this method is evaluated through scenarios of differing tether lengths, levels of observation error, and orbital eccentricities. Due to the need for the preliminary orbit determination to be achieved quickly, only short time intervals between observations were considered. The manner in which this preliminary orbit information can be used to obtain tether parameters for the subsequent differential correction process is also described.
Preliminary Orbit Determination of a Tethered Satellite Using the p-Iteration Method
NASA Astrophysics Data System (ADS)
Cicci, D. A.; Qualls, C.
2016-04-01
The possibility of the future deployment of tethered satellites has created a need for a preliminary orbit determination method capable of determining whether or not a satellite is tethered to another satellite. Classical preliminary orbit determination methods, which are used for untethered satellites, typically require two or more position vectors along with their respective observation times in order to determine a preliminary orbital element set. However, these conventional methods can't distinguish between an untethered satellite and a tethered one, whose motion is modified due to the presence of a tether force. The use of conventional methods on a satellite which is part of a tethered satellite system will result in the calculation of inaccurate orbital elements. Modifications have been made to the p-iteration preliminary orbit determination method in order to allow for the identification of these tethered satellites. The modifications allow for the calculation of a modified gravitational parameter, which can be used to distinguish between a tethered satellite and an untethered one. This paper applies this modified p-iteration method to the problem of the quick identification of a tethered satellite. The performance of this method is evaluated through scenarios of differing tether lengths, levels of observation error, and orbital eccentricities. Due to the need for the preliminary orbit determination to be achieved quickly, only short time intervals between observations were considered. The manner in which this preliminary orbit information can be used to obtain tether parameters for the subsequent differential correction process is also described.
Sub-meter GPS orbit determination and high precision user positioning - A demonstration
NASA Technical Reports Server (NTRS)
Lichten, Stephen M.; Bertiger, Willy I.; Katsigris, Eugenia C.
1988-01-01
High-accuracy orbit solutions have been obtained for GPS satellites, and submeter orbit accuracy is demonstrated for two well-tracked satellites. Orbit accuracy was tested based upon orbit repeatability from independent data sets, orbit prediction, ground baseline determination, and formal errors. Baselines of up to 2000 km in North America found with the GPS orbits show a daily repeatability of 0.3-1.5 parts in 10 to the 8th, and are found to agree well with VLBI solutions at the level of 0.3-3 parts in 10 to the 8th. Baselines were also determined between Florida and sites in the Caribbean region over 1000 km away, with a daily repeatability of 1-4 parts in 10 to the 8th.
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.
Observation error propagation on video meteor orbit determination
NASA Astrophysics Data System (ADS)
SonotaCo
2016-04-01
A new radiant direction error computation method on SonotaCo Network meteor observation data was tested. It uses single station observation error obtained by reference star measurement and trajectory linearity measurement on each video, as its source error value, and propagates this to the radiant and orbit parameter errors via the Monte Carlo simulation method. The resulting error values on a sample data set showed a reasonable error distribution that makes accuracy-based selecting feasible. A sample set of selected orbits obtained by this method revealed a sharper concentration of shower meteor radiants than we have ever seen before. The simultaneously observed meteor data sets published by the SonotaCo Network will be revised to include this error value on each record and will be publically available along with the computation program in near future.
Chou, Erh-Kang; Wu, Chao-I; Yu, Jack Chung-Kai; Chang, Sophia Chia-Ning
2009-05-01
Epistaxis is a frequent finding in patients with facial trauma. Herein, we report an unusual presentation of pediatric naso-orbital-ethmoid (NOE) fracture with epistaxis as the only initial symptom. The course of the patient's condition was later complicated by meningitis, related in part to the delay in diagnosis. A 3-year-old girl with preexisting upper respiratory symptoms was involved in a traffic accident, sustaining blunt trauma to the right side of her face. During the initial examination, only right-sided epistaxis was noted. Five days later, she developed febrile convulsion and was admitted to the intensive care unit with other signs of meningitis such as mental status change and neck stiffness. Her craniofacial computed tomographic scan showed a right-sided NOE fracture with minimal displacement and without dura tear. The cerebrospinal fluid culture grew Streptococcus pneumoniae, which may be due to ascending infection as a result of cribriform plate fracture. Intravenous antibiotic therapy was initiated with good response, and she was discharged from the hospital after 2 weeks. The presence of epistaxis and periorbital bruise, together with other symptoms and signs, helps in the identification of NOE and cribriform plate fracture. A high index of suspicion with repetitive computed tomographic scans is necessary to achieve correct early diagnosis. Parental antibiotic therapy is indicated if ascending cerebrospinal fluid infection develops. PMID:19461340
GPS orbit determination at the National Geodetic Survey
NASA Technical Reports Server (NTRS)
Schenewerk, Mark S.
1992-01-01
The National Geodetic Survey (NGS) independently generates precise ephemerides for all available Global Positioning System (GPS) satellites. Beginning in 1991, these ephemerides were produced from double-differenced phase observations solely from the Cooperative International GPS Network (CIGNET) tracking sites. The double-difference technique combines simultaneous observations of two satellites from two ground stations effectively eliminating satellite and ground receiver clock errors, and the Selective Availability (S/A) signal degradation currently in effect. CIGNET is a global GPS tracking network whose primary purpose is to provide data for orbit production. The CIGNET data are collected daily at NGS and are available to the public. Each ephemeris covers a single week and is available within one month after the data were taken. Verification is by baseline repeatability and direct comparison with other ephemerides. Typically, an ephemeris is accurate at a few parts in 10(exp 7). This corresponds to a 10 meter error in the reported satellite positions. NGS is actively investigating methods to improve the accuracy of its orbits, the ultimate goal being one part in 10(exp 8) or better. The ephemerides are generally available to the public through the Coast Guard GPS Information Center or directly from NGS through the Geodetic Information Service. An overview of the techniques and software used in orbit generation will be given, the current status of CIGNET will be described, and a summary of the ephemeris verification results will be presented.
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.
Orbit Determination of Non-cooperative Targets Using Laser Ranging Data at Changchun Station
NASA Astrophysics Data System (ADS)
Sun, J. N.; Liu, C. Z.; Fan, C. B.; Sun, M. G.
2015-09-01
The precise orbit determination software successfully processes the satellite laser ranging data of the non-cooperative targets in a single station. The insufficient observation data and the sole distribution data have become a principal difficulty in the orbit determination of the non-cooperative targets. Through the choices of dynamic models and the selections of solving parameters in the process of orbit determination, the condition equation can be solved with a convergence algorithm, and the orbit is obtained. The positional deviation obtained with the method of orbital overlap will be used as the accuracy index in calculating more groups of non-cooperative targets data. And the ranging deviation is obtained by comparing the trajectory information after orbit determination with the observation data uninvolved in orbit determination, which can be regarded as the externally coincident precision. The results show that the average ranging residual is 1.01 meters, the outer precision is 14.35 meters, and the precision of 1-day orbit prediction is 24.60 meters for non-cooperative target (4814).
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.
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.
Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)
NASA Technical Reports Server (NTRS)
Mesarch, Michael A.; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James
2007-01-01
This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF's orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.
Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)
NASA Technical Reports Server (NTRS)
Mesarch, Michael; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James
2007-01-01
This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF s orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.
Orbit Determination and Quality Assessment Using SELENE Tracking Data and Results
NASA Astrophysics Data System (ADS)
Goossens, Sander; Matsumoto, Koji; Ishihara, Yoshiaki; Liu, Qinghui; Kikuchi, Fuyuhiko; Noda, Hirotomo; Namiki, Noriyuki; Iwata, Takahiro
On September 14, 2007, the SELENE (KAGUYA) spacecraft were launched from Tanegashima Space Center in Japan. SELENE consists of three satellites: a main orbiter in a 100 km by 100 km circular, polar orbit, and two small subsatellites in 100 km by 2400 km (Rstar) and 100 km by 800 km (Vstar) elliptical, polar orbits. Until now, tracking of lunar satellites consisted of 2-way (or 3-way, where the upand downlink stations are different) tracking, leaving a gap in the tracking coverage over the far side of the Moon as the satellite cannot be tracked there from Earth. This severely hampers the determination of the global lunar gravity field, and, consequently, this also puts limits on the precision of orbits of lunar satellites. By employing 4-way Doppler tracking between the main orbiter and Rstar, the first direct tracking data of a satellite over the far side have been obtained, resulting in a newly determined global lunar gravity field. The existing 2-way tracking data set is furthermore complemented by differential VLBI tracking between Rstar and Vstar, providing a sensitivity perpendicular to the line-ofsight from station to satellite. This work focuses on aspects of orbit determination for the SELENE satellites, including the processing strategies for data types using multiple satellites. Orbit determination quality is described in terms of data fit and, where possible, orbit overlap statistics. For the main satellite, the on-board altimeter provides an independent check of orbit quality through crossovers, although they are not yet systematically included in the orbit determination process. The performance of the VLBI data in the orbit determination of the small subsatellites is also discussed. The newly determined global lunar gravity field models from SELENE are evaluated in several ways: their performance when used in orbit determination of previous lunar satellites, and their ability in orbit prediction. Covariance analysis shows the expected orbit quality
NASA Technical Reports Server (NTRS)
Morinelli, Patrick; Cosgrove, jennifer; Blizzard, Mike; Nicholson, Ann; Robertson, Mika
2007-01-01
This paper provides an overview of the launch and early orbit activities performed by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of five probes comprising the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft. The FDF was tasked to support THEMIS in a limited capacity providing backup orbit determination support for validation purposes for all five THEMIS probes during launch plus 30 days in coordination with University of California Berkeley Flight Dynamics Center (UCB/FDC). The FDF's orbit determination responsibilities were originally planned to be as a backup to the UCB/FDC for validation purposes only. However, various challenges early on in the mission and a Spacecraft Emergency declared thirty hours after launch placed the FDF team in the role of providing the orbit solutions that enabled contact with each of the probes and the eventual termination of the Spacecraft Emergency. This paper details the challenges and various techniques used by the GSFC FDF team to successfully perform orbit determination for all five THEMIS probes during the early mission. In addition, actual THEMIS orbit determination results are presented spanning the launch and early orbit mission phase. Lastly, this paper enumerates lessons learned from the THEMIS mission, as well as demonstrates the broad range of resources and capabilities within the FDF for supporting critical launch and early orbit navigation activities, especially challenging for constellation missions.
NASA Technical Reports Server (NTRS)
Morinelli, Patrick; Cosgrove, Jennifer; Blizzard, Mike; Robertson, Mike
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)2. The FDF's orbit determination responsibilities were originally planned to be as a backup to the UCB/FDC for validation purposes only. However, various challenges early on in the mission and a Spacecraft Emergency declared thirty hours after launch placed the FDF team in the role of providing the orbit solutions that enabled contact with each of the probes and the eventual termination of the Spacecraft Emergency. This paper details the challenges and various techniques used by the GSFC FDF team to successfully perform orbit determination for all five THEMIS probes during the early mission. In addition, actual THEMIS orbit determination results are presented spanning the launch and early orbit mission phase. Lastly, this paper enumerates lessons learned from the THEMIS mission, as well as demonstrates the broad range of resources and capabilities within the FDF for supporting critical launch and early orbit navigation activities, especially challenging for constellation missions.
NASA Technical Reports Server (NTRS)
Head, D. E.; Mitchell, K. L.
1967-01-01
Program computes the thermal environment of a spacecraft in a lunar orbit. The quantities determined include the incident flux /solar and lunar emitted radiation/, total radiation absorbed by a surface, and the resulting surface temperature as a function of time and orbital position.
NASA Astrophysics Data System (ADS)
Shefer, V. A.
2014-12-01
constructed using the A1 and A2 algorithms (2-4 orders of magnitude in coordinates and 4-7 orders of magnitude in velocities higher) compared to the accuracy of the approximation by Keplerian orbits with decreasing the reference arc of the trajectory. Here, the higher is the efficiency of the algorithms A1 and A2, the smaller are the values of the topocentric distances, i.e., the greater are the perturbations caused by the Earth's gravitation. The advantage of Algorithm A2 over Algorithm A1 in accuracy extends approximately one order of magnitude. The minimal methodic errors of the position vector by using the A1 and A2 algorithms range from several meters in the case of the asteroid Apophis to several millimeters in the case of the asteroid 2012 DA14. Hence, the numerical examples analyzed in this work lead us to conclude that the proposed in [1, 2] methods for determination of an intermediate perturbed orbit from range and range rate measurements at three time points allow for significantly raising the accuracy of the calculation of the initial asteroid orbits in comparison with the algorithm based on the finding the unperturbed Keplerian orbit. The shorter is the orbital arc specified by the extreme time points, the greater is the advantage of the algorithms suggested over the algorithms of the traditional approach in the accuracy. The advantage of the algorithms suggested in the accuracy increases with raising the perturbations too, which is especially important for calculation of the initial trajectories of the space objects detected in the Earth's neighbourhood. The work was supported by the Ministry of Education and Science of the Russian Federation, project no. 2014/223(1567).
Laser ranging network performance and routine orbit determination at D-PAF
NASA Technical Reports Server (NTRS)
Massmann, Franz-Heinrich; Reigber, C.; Li, H.; Koenig, Rolf; Raimondo, J. C.; Rajasenan, C.; Vei, M.
1993-01-01
ERS-1 is now about 8 months in orbit and has been tracked by the global laser network from the very beginning of the mission. The German processing and archiving facility for ERS-1 (D-PAF) is coordinating and supporting the network and performing the different routine orbit determination tasks. This paper presents details about the global network status, the communication to D-PAF and the tracking data and orbit processing system at D-PAF. The quality of the preliminary and precise orbits are shown and some problem areas are identified.
NASA Technical Reports Server (NTRS)
Fang, B. T.
1979-01-01
The TDRS was modeled as a combination of a sun-pointing solar panel and earth-pointing plate. Based on this model, explanations are given for the following orbit determination error characteristics: inherent limits in orbital accuracy, the variation of solar pressure induced orbital error with time of the day of epoch, the insensitivity of range-rate orbits to GM error, and optimum bilateration baseline.
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.
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.
Impact of tracking station distribution structure on BeiDou satellite orbit determination
NASA Astrophysics Data System (ADS)
Zhang, Rui; Zhang, Qin; Huang, Guanwen; Wang, Le; Qu, Wei
2015-11-01
The racking station distribution structure plays an important role in GNSS satellite orbit determination. Due to the current satellite distribution of the BeiDou satellite navigation system (BDS), the problem how to construct a reasonable distribution of tracking stations to obtain BDS satellite orbits with high precision has become a highly imperative issue. Based on the theory of dynamic orbit determination, two different station distributions were analyzed to study their impact on BDS precise and real-time orbit determination. Subsequently, the impact of Satellite Position Dilution of Precision (SPDOP) values on orbit determination was analyzed. Finally, an improved scheme for the tracking station distribution was designed based on the original scheme. The numerical results show that the SPDOP value can be used to evaluate the contribution of the tracking stations distribution on the BDS IGSO and MEO satellites orbit determination. In addition, the tracking stations which focus on the Asia-Pacific region play a key role in current BDS orbit determination.
Precise Orbit Determination for the GEOSAT Follow-On Spacecraft
NASA Technical Reports Server (NTRS)
Lemoine, Frank G.; Rowlands, David D.; Zelensky, Nikita P.; Luthcke, Scott B.; Cox, Christopher M.; Marr, Gregory C.
1999-01-01
The US Navy's GEOSAT Follow-On spacecraft was launched on February 10, 1998 with its primary mission objective to map the oceans using a radar altimeter. The spacecraft tracking complement consists of GPS receivers, a laser retroreflector and Doppler beacons. Since the GPS receivers have not yet returned reliable data, the only means of providing high-quality precise orbits has been though satellite laser ranging (SLR). SLR has tracked the spacecraft since April 22, 1998, and an average of 7 passes per day have been obtained from US and foreign stations. Since the predicted radial orbit error due to the gravity field is only two to three cm, the largest contributor to the high SLR residuals (10 cm) is the mismodelling of the non-conservative forces. The SLR residuals show a clear correlation with beta prime (solar elevation) angle, peaking in mid-August 1998 when the beta prime angle reached -80 to -90 degrees. We report in this paper on the analysis of the GFO tracking data (SLR, Doppler, and if available GPS) using GEODYN, and on the tuning of the non-conservative force model and the gravity model using these data.
Orbit determination with the two-body integrals
NASA Astrophysics Data System (ADS)
Gronchi, G. F.; Dimare, L.; Milani, A.
2010-07-01
We investigate a method to compute a finite set of preliminary orbits for solar system bodies using the first integrals of the Kepler problem. This method is thought for the applications to the modern sets of astrometric observations, where often the information contained in the observations allows only to compute, by interpolation, two angular positions of the observed body and their time derivatives at a given epoch; we call this set of data attributable. Given two attributables of the same body at two different epochs we can use the energy and angular momentum integrals of the two-body problem to write a system of polynomial equations for the topocentric distance and the radial velocity at the two epochs. We define two different algorithms for the computation of the solutions, based on different ways to perform elimination of variables and obtain a univariate polynomial. Moreover we use the redundancy of the data to test the hypothesis that two attributables belong to the same body ( linkage problem). It is also possible to compute a covariance matrix, describing the uncertainty of the preliminary orbits which results from the observation error statistics. The performance of this method has been investigated by using a large set of simulated observations of the Pan-STARRS project.
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.
Precise orbit determination for the FORMOSAT-3/COSMIC satellite mission using GPS
NASA Astrophysics Data System (ADS)
Hwang, Cheinway; Tseng, Tzu-Pang; Lin, Tingjung; Švehla, Dražen; Schreiner, Bill
2009-05-01
The joint Taiwan-US mission FORMOSAT-3/ COSMIC (COSMIC) was launched on April 17, 2006. Each of the six satellites is equipped with two POD antennas. The orbits of the six satellites are determined from GPS data using zero-difference carrier-phase measurements by the reduced dynamic and kinematic methods. The effects of satellite center of mass (COM) variation, satellite attitude, GPS antenna phase center variation (PCV), and cable delay difference on the COSMIC orbit determination are studied. Nominal attitudes estimated from satellite state vectors deliver a better orbit accuracy when compared to observed attitude. Numerical tests show that the COSMIC COM must be precisely calibrated in order not to corrupt orbit determination. Based on the analyses of the 5 and 6-h orbit overlaps of two 30-h arcs, orbit accuracies from the reduced dynamic and kinematic solutions are nearly identical and are at the 2-3 cm level. The mean RMS difference between the orbits from this paper and those from UCAR (near real-time) and WHU (post-processed) is about 10 cm, which is largely due to different uses of GPS ephemerides, high-rate GPS clocks and force models. The kinematic orbits of COSMIC are expected to be used for recovery of temporal variations in the gravity field.
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
NASA Technical Reports Server (NTRS)
Kennedy, Brian; Abrahamson, Matt; Ardito, Alessandro; Han, Dongsuk; Haw, Robert; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew
2013-01-01
The Dawn spacecraft was launched on September 27th, 2007. Its mission is to consecutively rendezvous with and observe the two largest bodies in the asteroid belt, Vesta and Ceres. It has already completed over a year's worth of direct observations of Vesta (spanning from early 2011 through late 2012) and is currently on a cruise trajectory to Ceres, where it will begin scientific observations in mid-2015. Achieving this data collection required careful planning and execution from all spacecraft teams. Dawn's Orbit Determination (OD) team was tasked with accurately predicting the trajectory of the Dawn spacecraft during the Vesta science phases, and also determining the parameters of Vesta to support future science orbit design. The future orbits included the upcoming science phase orbits as well as the transfer orbits between science phases. In all, five science phases were executed at Vesta, and this paper will describe some of the OD team contributions to the planning and execution of those phases.
Initial Determinations of Ionospheric Electric Fields and Joule Heating from MAVEN Observations
NASA Astrophysics Data System (ADS)
Fillingim, M. O.; Fogle, A. L.; Aleryani, O.; Dunn, P.; Lillis, R. J.; McFadden, J. P.; Connerney, J. E. P.; Mahaffy, P. R.; Andersson, L.; Ergun, R.
2015-12-01
MAVEN provides in-situ measurements of the neutral and ion species as well as the magnetic field throughout the ionosphere of Mars. By combining these measurements, we are able to calculate both the ionospheric currents and the ionospheric conductivity. It is then straightforward to determine the electric field in the collisional ionosphere from a simplified Ohm's law. In addition, we can also estimate the amount of Joule heating in the ionosphere from j · E. Here, we show initial determinations of both ionospheric electric fields and Joule heating using MAVEN data. The electric fields are highly variable from orbit-to-orbit suggesting that the ionospheric electrodynamics can change on timescales of several hours. These changes may be driven by changes in the upstream solar wind and IMF or may result from dynamical variations of thermospheric neutral winds.
Orbiting Deep Space Relay Station (ODSRS). Volume 1: Requirement determination
NASA Technical Reports Server (NTRS)
Hunter, J. A.
1979-01-01
The deep space communications requirements of the post-1985 time frame are described and the orbiting deep space relay station (ODSRS) is presented as an option for meeting these requirements. Under current conditions, the ODSRS is not yet cost competitive with Earth based stations to increase DSN telemetry performance, but has significant advantages over a ground station, and these are sufficient to maintain it as a future option. These advantages include: the ability to track a spacecraft 24 hours per day with ground stations located only in the USA; the ability to operate at higher frequencies that would be attenuated by Earth's atmosphere; and the potential for building very large structures without the constraints of Earth's gravity.
The use of mixed observations from one station to determine the preliminary orbits
NASA Astrophysics Data System (ADS)
Baghos, B. B.
The use of mixed observations from a laser tracking station whose geodetic coordinates are known, to determine a preliminary orbit is examined. The method has the advantage of being able to determine an osculating orbit without using iterative processes or the truncated f and g series. By using mixed data obtained from Ondrejov Observatory for the satellite, Starlette, the numerical stability of the method is discussed. A computer routine in basic FORTRAN, which illlustrates the procedure, is given.
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
42 CFR 405.924 - Actions that are initial determinations.
Code of Federal Regulations, 2014 CFR
2014-10-01
... 42 Public Health 2 2014-10-01 2014-10-01 false Actions that are initial determinations. 405.924 Section 405.924 Public Health CENTERS FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH AND HUMAN SERVICES MEDICARE PROGRAM FEDERAL HEALTH INSURANCE FOR THE AGED AND DISABLED Determinations, Redeterminations, Reconsiderations, and Appeals...
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.
Determinants of Cigarette Smoking Initiation in Jordanian Schoolchildren: Longitudinal Analysis
Attonito, Jennifer; Madhivanan, Purnima; Yi, Qilong; Mzayek, Fawaz; Maziak, Wasim
2015-01-01
Objective: To identify determinants of cigarette smoking initiation, by gender, among schoolchildren in Irbid, Jordan. Methods: Between 2008 and 2011, data were collected annually using self-reported questionnaires over 4-years in a prospective cohort of 1,781 students recruited from all 7th grade classes in 19 secondary schools, selected out of a total 60, using probability-proportionate-to-size method. Independent predictors of smoking initiation were identified among the cigarette naïve participants (N = 1,454) with mixed-effect multivariable logistic regression. Results: Participants were 12.6 years of age on average at baseline. 29.8% of the 1,454 students (37.2% of boys and 23.7% of girls) initiated cigarette smoking by 10th grade. Of those who initiated (n = 498), 47.2% of boys and 37.2% of girls initiated smoking in the 8th grade. Determinants of cigarette smoking initiation included ever smoking a waterpipe, low cigarette refusal self-efficacy, intention to start smoking cigarettes, and having friends who smoked. For girls, familial smoking was also predictive of cigarette initiation. Conclusion: This study shows that many Jordanian youth have an intention to initiate cigarette smoking and are susceptible to cigarette smoking modeled by peers and that girls are influenced as well by familial cigarette smoking. Prevention efforts should be tailored to address culturally relevant gender norms, help strengthen adolescents’ self-efficacy to refuse cigarettes, and foster strong non-smoking social norms. PMID:25143297
Taking advantage of the MEMO orbiter to improve the determination of Mars' gravity field.
NASA Astrophysics Data System (ADS)
Rosenblatt, P.; Le Maitre, S.; Marty, J. C.; Duron, J.; Dehant, V.
2007-08-01
In the context of future ESA's mission to Mars, it is proposed an orbiter named MEMO (Mars Escape and Magnetic Orbiter) to especially improve the measurement of the atmospheric escape and the magnetic field of the planet. Its orbit is planned to have an inclination of 77 degrees and periapsis and apoapsis altitude of 130 km and 1000 km, respectively. In addition, such an orbit is scheduled to be maintained during one Martian year. This differs from the usual near-polar, near-circular orbit with a periapsis altitude of at least 200 km, such as for Mars Reconnaissance Orbiter (MRO). Even if the MEMO orbiter is not dedicated to Mars' gravity field investigation, we propose to take this opportunity to improve our knowledge of Mars' gravity field. Indeed, the sensitivity of an orbiter to the gravity field strongly depends on the semi-major axis, inclination and eccentricity of its orbit. In this study, we quantitatively estimate the improvement on the determination of local gravity anomalies, of seasonal variations of the first zonal harmonics and of the k2 Love number of Mars. We base our work on both analytical and numerical approaches in order to simulate the Mars' gravity field determination from spacecraft tracking data from the Earth.We also add in our simulations the possibility to have an accelerometer onboard the MEMO spacecraft. Indeed, if it is placed at the center of mass of the spacecraft, it could provide measurements of the non-gravitational forces acting on it, especially the atmospheric drag. A good determination of the contribution of this force to the spacecraft motion would bring information about the atmospheric density at altitude between 100 and 200 km, and would improve the gravity field determination from tracking data of the spacecraft.
Ren Shulin; Fu Yanning E-mail: fyn@pmo.ac.c
2010-05-15
Untill now, the Hipparcos intermediate astrometric data (HIAD) have contributed little to the full orbit determination of double-lined spectroscopic binaries (SB2s). This is because the photocenter of such a binary system is usually not far from the system mass center, and its orbital wobble is generally weak with respect to the accuracy of the HIAD. However, the HIAD have been recently revised and the accuracy is increased by a factor of 2.2 in the total weight. Therefore, it is interesting to see if the revised HIAD can be used in the orbit determination at least for some SB2s. In this paper, we first search the 9th Catalogue of Orbits of Spectroscopic Binaries (S{sub B{sup 9}}) for SB2s with reliable spectroscopic orbital solutions and with periods between 50 days and 3.2 years. This leaves us with 56 systems. The full orbital solutions of these systems are then determined from the HIAD by a highly efficient grid search method developed in this paper. The high efficiency is achieved by reducing the number of nonlinear model parameters to one, and by allowing all parameters to be adjustable within a region centered at each grid point. After a variety of tests, we finally accept orbital solutions of 13 systems. Among these systems, six (HIP 677, 20894, 87895, 95995, 101382, and 111170) are well resolved with reliable interferometric data. Orbital solutions from these data are consistent with our results. The full orbital solutions of the other seven systems (HIP 9121, 17732, 32040, 57029, 76006, 102431, and 116360) are determined for the first time.
TDRSS-user orbit determination using batch least-squares and sequential methods
NASA Technical Reports Server (NTRS)
Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.
1993-01-01
The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.
Radial orbit error reduction and sea surface topography determination using satellite altimetry
NASA Technical Reports Server (NTRS)
Engelis, Theodossios
1987-01-01
A method is presented in satellite altimetry that attempts to simultaneously determine the geoid and sea surface topography with minimum wavelengths of about 500 km and to reduce the radial orbit error caused by geopotential errors. The modeling of the radial orbit error is made using the linearized Lagrangian perturbation theory. Secular and second order effects are also included. After a rather extensive validation of the linearized equations, alternative expressions of the radial orbit error are derived. Numerical estimates for the radial orbit error and geoid undulation error are computed using the differences of two geopotential models as potential coefficient errors, for a SEASAT orbit. To provide statistical estimates of the radial distances and the geoid, a covariance propagation is made based on the full geopotential covariance. Accuracy estimates for the SEASAT orbits are given which agree quite well with already published results. Observation equations are develped using sea surface heights and crossover discrepancies as observables. A minimum variance solution with prior information provides estimates of parameters representing the sea surface topography and corrections to the gravity field that is used for the orbit generation. The simulation results show that the method can be used to effectively reduce the radial orbit error and recover the sea surface topography.
OREMUS: Prototype of an orbit determination center dedicated to mini-satellite control
NASA Astrophysics Data System (ADS)
Laporte, Francois; Campan, Genevieve; Conessa, Huguette
1993-01-01
The OREMUS project aims to achieve at moderate cost a ground segment prototype for the purposes of orbiting satellite operations. In order to demonstrate the feasibility of a general system capable of satisfying the requirements of various users, main functions to be improved are data acquisition, orbit restitution, orbital maneuver calculations, and orbit linked expected operations. MERCATOR system architecture, components, capabilities, interest, and performance are recalled as well as the OOC (Operational Orbit determination Center) needs as regards the software. Characteristics and advantages of a new man machine interface called SISSI are presented. The OREMUS general system is based on MERCATOR and SISSI functions, and on the recovery of former modulus used in spatial mechanics. Descriptor files are documented by the users. Prototype main functions, data acquisition, orbit restitution, attitude restitution, orbital maneuver calculations, orbit linked expected operations, file management, and parameters graphic visualization are available in a control ground segment. These functions are performed by means of modulus; the linking and interfaces of which are managed by the man machine interface.
Orbit Determination from Combined Radar and Optical Tracks during XMM Contingency Operations
NASA Astrophysics Data System (ADS)
Flohrer, T.; Krag, H.; Klinkrad, H.; Kuusela, J.; Leushacke, L.; Schildknecht, T.; Ploner, M.
2009-03-01
On 18 October 2008 the operators of ESA's X-ray Multi- Mirror Mission (XMM-Newton) lost contact with the satellite. XMM is one of Europe's largest scientific satellites. It resides in a highly eccentric (21700 km x 99500 km) orbit with an inclination of 58 deg. The XMM operators asked to support the analysis of the contingency situation, in particular to acquire tracking data of the noncooperative target via suitable tracking facilities, and to determine a precise orbit. Any information on orbital states and attitude was highly desirable in order to better understand the situation and to ensure proper followups with the ground facilities during communication attempts. We present the fusion of radar and optical observations into a common orbit determination of a noncooperative target using the predicted orbit as a-priori information. Three European sensors participated in the adhoc tracking campaign: the Tracking and Imaging Radar (TIRA) of the Forschungsgesellschaft für Angewandte Naturwissenschaften (FGAN) near Bonn, Germany, the ESA Space Debris telescope at Tenerife, Spain, and the telescopes ZIMLAT and ZimSMART at the Zimmerwald observatory of the Astronomical Institute of the University of Bern (AIUB) in Switzerland. All sensors were able to observe XMM close to the predicted positions. In the meantime the New Norcia ground station could establish a weak carrier-link. This finally led to re-establishing full radio contact. We validate the quality of the orbit determination through a comparison with the operational orbit. This work demonstrates the generation of orbit information for passive bodies by using European sensors only, even if the orbit is highly eccentric.
A demonstration of high precision GPS orbit determination for geodetic applications
NASA Technical Reports Server (NTRS)
Lichten, S. M.; Border, J. S.
1987-01-01
High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.
Investigations to Determine the Origin of the Solar Wind with SPICE and SolarOrbiter
NASA Astrophysics Data System (ADS)
Hassler, Donald M.; DeForest, C.; Wilkinson, E.; Davila, J.; SPICE Team
2011-05-01
At large spatial scales, the structure of the solar wind and it's mapping back to the solar corona, is thought to be reasonably well understood. However, the detailed structure of the various source regions at chromospheric and transition region heights is extremely complex, and less well understood. Determining this connection between heliospheric structures and their source regions at the Sun is one of the overarching objective of the Solar Orbiter mission. During perihelion segments of its orbit, when the spacecraft is in quasi-corotation with the Sun, Solar Orbiter will determine the plasma parameters and compositional signatures of the solar wind, which can be compared directly with the spectroscopic signatures of coronal ions with differing charge-to-mass ratios and FIP. One of the key instruments on the Solar Orbiter mission to make these remote sensing measurements is the SPICE (Spectral Imaging of the Coronal Environment) imaging spectrograph. SPICE will provide the images and plasma diagnostics needed to characterize the plasma state in different source regions, from active regions to quiet Sun to coronal holes. By comparing composition, plasma parameters, and low/high FIP ratios of structures remotely, with those measured directly at the Solar Orbiter spacecraft, Solar Orbiter will provide the first direct link between solar wind structures and their source regions at the Sun. This talk will provide a background of previous compositional correlation measurements and an outline of the method to be used for comparing the spectroscopic and in-situ plasma parameters to be measured with Solar Orbiter.
Orbit determination strategy and results for the Pioneer 10 Jupiter mission
NASA Technical Reports Server (NTRS)
Wong, S. K.; Lubeley, A. J.
1974-01-01
Pioneer 10 is the first earth-based vehicle to encounter Jupiter and occult its moon, Io. In contributing to the success of the mission, the Orbit Determination Group evaluated the effects of the dominant error sources on the spacecraft's computed orbit and devised an encounter strategy minimizing the effects of these error sources. The encounter results indicated that: (1) errors in the satellite model played a very important role in the accuracy of the computed orbit, (2) encounter strategy was sound, (3) all mission objectives were met, and (4) Jupiter-Saturn mission for Pioneer 11 is within the navigation capability.
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.
NASA Astrophysics Data System (ADS)
Asada, Hideki
2006-11-01
There exists a very classical inverse problem regarding orbit determination of a binary system: "when an orbital plane of two bodies is inclined with respect to the line of sight, observables are their positions projected onto a celestial sphere. How do we determine the orbital elements from observations?" A "complete exact solution" has been found. It is reviewed with some related topics.
TerraSAR-X precise orbit determination with real-time GPS ephemerides
NASA Astrophysics Data System (ADS)
Wermuth, M.; Hauschild, A.; Montenbruck, O.; Kahle, R.
2012-09-01
For active and future Earth observation missions, the availability of near real-time precise orbit information is becoming more and more important. The latency and quality of precise orbit determination results is mainly driven by the availability of precise GPS ephemerides and clocks. In order to have high-quality GPS ephemerides and clocks available at real-time, the German Space Operations Center (GSOC) has developed the real-time clock estimation system RETICLE. The system receives data streams with GNSS observations from the global tracking network of the International GNSS Service (IGS) in real-time. Using the known station position, RETICLE estimates precise GPS satellite clock offsets and drifts based on the most recent available ultra rapid predicted orbits provided by the IGS. The clock offset estimates have an accuracy of better than 0.3 ns and are globally valid. The latency of the estimated clocks is approximately 7 s after the observation epoch. Another limiting factor is the frequency of satellite downlinks and the latency of the data transfer from the ground station to the operations center. Therefore a near real-time scenario using GPS observation data from the TerraSAR-X mission is examined in which the satellite has about one ground station contact per orbit or respectively one contact in 90 min. This test campaign shows that precise orbits can be obtained in near real-time. With the use of estimated clocks an orbit accuracy of better than 10 cm (3D-RMS) can be obtained. The evaluation of satellite laser ranging (SLR) observations shows residuals of 2.1 cm (RMS) for orbits using RECTICLE and residuals of 4.2 cm (RMS) for orbits using the IGS ultra rapid ephemerides and clocks products. Hence the use of estimated clocks improves the orbit determination accuracy significantly (˜factor 2) compared to using predicted clocks.
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.
Orbit Determination for CE-2 Libration Flight and Asteroid Exploration Trial
NASA Astrophysics Data System (ADS)
Cao, J. F.
2016-01-01
Setting within the context of the flight trial of CE-2 (Chang'e 2) around the Sun-terrestrial libration point, the asteroid exploration as well as the YH-1 Mars exploration mission, this paper conducted various related studies on orbit determination techniques for deep space exploration. The research results provided high-precision orbit support for the successful photographing of the Toutatis. This paper also carried out preliminary orbit determination studies on YH-1 mission. Although the study findings can not be used directly in the Mars exploration mission, they can still be useful for the future explorations. This thesis is composed of the following five aspects. (1)Reviewed the statistical orbit determination theory, and gave a description of the spatiotemporal frame of reference, dynamical model issues, methods of estimation, perturbation analysis theory, as well as the algorithms for considering covariance analysis. (2)Developed the observational model for the deep space exploration. Based on theoretical analysis, the models of ranging, ranging rate, and VLBI (Very Long Baseline Interferometry) are derived. During the modeling process, the algorithm is optimized to improve the computational efficiency without deteriorating the accuracy. In addition, with the spin-stabilized characteristic of CE-2 in its cruise phase taken into consideration, a spin stabilization correction model of the tracking data is constructed, which not only meets the requirement of data correction, but also can estimate the alignment of antenna. (3)Carried out a study on the selection of integration center for CE-2 libration flight trial. The result shows that the Earth is most suitable for orbital prediction. A precise satellite ephemeris for CE-2's flight trial is provided. The transformation relation between the spatial-fixed coordinate system and the rotation coordinate system is constructed. An orbital accuracy of 2--10 km in the whole flight process and 5 km for the stable
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.
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.
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.
Evaluation the initial estimators using deterministic minimum covariance determinant algorithm
NASA Astrophysics Data System (ADS)
Alrawashdeh, Mufda Jameel; Sabri, Shamsul Rijal Muhammad; Ismail, Mohd Tahir
2014-07-01
The aim of the study is to examine five initial estimators introduced by Hubert et al. [1] with five additional new initial estimators by using the Deterministic Minimum Covariance Determinant algorithm, DetMCD. The objective of the DetMCD is to robustify the location and scatter matrix parameters. Since these parameters are highly influenced by the presence of outliers, the DetMCD is a newly highly robust algorithm, where it is constructed to overcome the outlier's problem. DetMCD precedes the non-random subsets, which computes a small number of deterministic initial estimators and followed by concentration steps. Here, we are going to compare the DetMCD algorithm based on two groups of estimators - one with original five Huberts' estimators and the other five new estimators. The determinant values of these estimators are observed to evaluate the performance via several cases.
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.
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2012 CFR
2012-01-01
... 10 Energy 1 2012-01-01 2012-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations § 9.29 Appeal... for Investigations, the appeal must be in writing directed to the Inspector General and sent to...
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 10 Energy 1 2010-01-01 2010-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations § 9.29 Appeal... for Investigations, the appeal must be in writing directed to the Inspector General and sent to...
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2014 CFR
2014-01-01
... 10 Energy 1 2014-01-01 2014-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations § 9.29 Appeal... for Investigations, the appeal must be in writing directed to the Inspector General and sent to...
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2013 CFR
2013-01-01
... 10 Energy 1 2013-01-01 2013-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations § 9.29 Appeal... for Investigations, the appeal must be in writing directed to the Inspector General and sent to...
10 CFR 9.29 - Appeal from initial determination.
Code of Federal Regulations, 2011 CFR
2011-01-01
... 10 Energy 1 2011-01-01 2011-01-01 false Appeal from initial determination. 9.29 Section 9.29 Energy NUCLEAR REGULATORY COMMISSION PUBLIC RECORDS Freedom of Information Act Regulations § 9.29 Appeal... for Investigations, the appeal must be in writing directed to the Inspector General and sent to...
42 CFR 476.83 - Initial denial determinations.
Code of Federal Regulations, 2012 CFR
2012-10-01
... 42 Public Health 4 2012-10-01 2012-10-01 false Initial denial determinations. 476.83 Section 476.83 Public Health CENTERS FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) QUALITY IMPROVEMENT ORGANIZATIONS UTILIZATION AND QUALITY CONTROL REVIEW Review Responsibilities of Utilization and Quality Control...
42 CFR 476.83 - Initial denial determinations.
Code of Federal Regulations, 2013 CFR
2013-10-01
... 42 Public Health 4 2013-10-01 2013-10-01 false Initial denial determinations. 476.83 Section 476.83 Public Health CENTERS FOR MEDICARE & MEDICAID SERVICES, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) QUALITY IMPROVEMENT ORGANIZATIONS UTILIZATION AND QUALITY CONTROL REVIEW Review Responsibilities of Utilization and Quality Control...
24 CFR 599.301 - Initial determination of threshold requirements.
Code of Federal Regulations, 2014 CFR
2014-04-01
... HOUSING AND URBAN DEVELOPMENT COMMUNITY FACILITIES RENEWAL COMMUNITIES Evaluation of Applications... 24 Housing and Urban Development 3 2014-04-01 2013-04-01 true Initial determination of threshold requirements. 599.301 Section 599.301 Housing and Urban Development Regulations Relating to Housing and...
24 CFR 599.301 - Initial determination of threshold requirements.
Code of Federal Regulations, 2010 CFR
2010-04-01
... HOUSING AND URBAN DEVELOPMENT COMMUNITY FACILITIES RENEWAL COMMUNITIES Evaluation of Applications... 24 Housing and Urban Development 3 2010-04-01 2010-04-01 false Initial determination of threshold requirements. 599.301 Section 599.301 Housing and Urban Development Regulations Relating to Housing and...
24 CFR 599.301 - Initial determination of threshold requirements.
Code of Federal Regulations, 2011 CFR
2011-04-01
... HOUSING AND URBAN DEVELOPMENT COMMUNITY FACILITIES RENEWAL COMMUNITIES Evaluation of Applications... 24 Housing and Urban Development 3 2011-04-01 2010-04-01 true Initial determination of threshold requirements. 599.301 Section 599.301 Housing and Urban Development Regulations Relating to Housing and...
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
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
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
Goetz, A.F.H.; Rowan, L.C.; Kingston, M.J.
1982-01-01
A shuttle-borne radiometer containing ten channels in the reflective infrared has demonstrated that direct identification of carbonates and hydroxyl-bearing minerals is possible by remote measurement from Earth orbit. Copyright ?? 1982 AAAS.
GOCE Precise Orbit Determination for the Entire Mission- Challenges in the Final Mission Phase
NASA Astrophysics Data System (ADS)
Jaggi, A.; Bock, H.; Meyer, U.
2015-03-01
The Gravity field and steady-state Ocean Circulation Explorer (GOCE), ESA’s first Earth Explorer core mission, was launched on March 17, 2009 into a sun-synchronous dusk-dawn orbit and eventually re-entered into the Earth’s atmosphere on November 11, 2013. A precise science orbit (PSO) product was provided by the GOCE High-level Processing Facility (HPF) from the GPS high-low Satellite-to-Satellite Tracking (hl-SST) data from the beginning until the very last days of the mission. We recapitulate the PSO procedure and refer to the results achieved until the official end of the GOCE mission on October 21, 2013, where independent validations with Satellite Laser Ranging (SLR) measurements confirmed a high quality of the PSO product of about 2 cm 1-D RMS. We then focus on the period after the official end of the mission, where orbits could still be determined thanks to the continuously running GPS receivers delivering high quality data until a few hours before the re-entry into the Earth’s atmosphere. We address the challenges encountered for orbit determination during these last days and report on adaptions in the PSO procedure to also obtain good orbit results at the unprecedented low orbital altitudes below 224 km.
NASA Technical Reports Server (NTRS)
Peters, Palmer N.; Gregory, John C.
1992-01-01
Images produced by pinhole cameras using film sensitive to atomic oxygen provide information on the ratio of spacecraft orbital velocity to the most probable thermal speed of oxygen atoms, provided the spacecraft orientation is maintained stable relative to the orbital direction. Alternatively, information on the spacecraft attitude relative to the orbital velocity can be obtained, provided that corrections are properly made for thermal spreading and a corotating atmosphere. The Long Duration Exposure Facility (LDEF) orientation, uncorrected for a corotating atmosphere, was determined to be yawed 8.0 +/- 0.4 degrees from its nominal attitude, with an estimated +/- 0.35 degree oscillation in yaw. The integrated effect of inclined orbit and corotating atmosphere produces an apparent oscillation in the observed yaw direction, suggesting that the LDEF attitude measurement will indicate even better stability when corrected for a corotating atmosphere. The measured thermal spreading is consistent with major exposure occurring during high solar activity, which occurred late during the LDEF mission.
Interplanetary Departure Stage Navigation by Means of Liaison Orbit Determination Architecture
NASA Technical Reports Server (NTRS)
McGranaghan, Ryan M.; Leonard, Jason M.; Fujimoto, Kohei; Parker, Jeffrey S.; Anderson, Rodney L.; Born, George H.
2013-01-01
Autonomous orbit determination for departure stages of interplanetary trajectories is conducted by means of realistic radiometric observations between the departing spacecraft and a satellite orbiting the first lunar libration point. Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) is used to estimate the orbit solution. This paper uses high-fidelity simulations to explore the utilization of LiAISON in providing improved accuracy for interplanetary departure missions. The use of autonomous navigation to supplement current techniques for interplanetary spacecraft is assessed using comparisons with groundbased navigation. Results from simulations including the Mars Science Laboratory, Mars Exploration Rover, and Cassini are presented. It is shown that observations from a dedicated LiAISON navigation satellite could be used to supplement ground-based measurements and significantly improve tracking performance.
DETERMINATION OF ORBITAL ELEMENTS OF SPECTROSCOPIC BINARIES USING HIGH-DISPERSION SPECTROSCOPY
Katoh, Noriyuki; Itoh, Yoichi; Toyota, Eri; Sato, Bun'ei
2013-02-01
Orbital elements of 37 single-lined spectroscopic binary systems (SB1s) and 5 double-lined spectroscopic binary systems (SB2s) were determined using high-dispersion spectroscopy. To determine the orbital elements accurately, we carried out precise Doppler shift measurements using the HIgh Dispersion Echelle Spectrograph mounted on the Okayama Astrophysical Observatory 1.88 m telescope. We achieved a radial-velocity precision of {approx}10 m s{sup -1} over seven years of observations. The targeted binaries have spectral types between F5 and K3, and are brighter than the 7th magnitude in the V band. The orbital elements of 28 SB1s and 5 SB2s were determined at least 10 times more precisely than previous measurements. Among the remaining nine SB1s, five objects were found to be single stars, and the orbital elements of four objects were not determined because our observations did not cover the entire orbital period. We checked the absorption lines from the secondary star for 28 SB1s and found that three objects were in fact SB2s.
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.
Cassini Orbit Determination Performance during Saturn Satellite Tour: August 2005 - January 2006
NASA Technical Reports Server (NTRS)
Antreasian, Peter G.; Bordi, J. J.; Criddle, K. E.; Ionasescu, R.; Jacobson, R. A.; Jones, J. B.; MacKenzie, R. A.; Parcher, D. W.; Pelletier, F. J.; Roth, D. C.; Stauch, J. R.
2007-01-01
During the period spanning the second Enceladus flyby in July 2005 through the eleventh Titan encounter in January 2006, the Cassini spacecraft was successfully navigated through eight close-targeted satellite encounters. Three of these encounters included the 500 km flybys of the icy satellites Hyperion, Dione and Rhea and five targeted flybys of Saturn's largest moon, Titan. This paper will show how our refinements to Saturn's satellite ephemerides have improved orbit determination predictions. These refinements include the mass estimates of Saturn and its satellites by better than 0.5%. Also, it will be shown how this better orbit determination performance has helped to eliminate several statistical maneuvers that were scheduled to clean-up orbit determination and/or maneuver-execution errors.
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
NASA Astrophysics Data System (ADS)
Stassun, Keivan
2016-05-01
We summarize the current state-of-the-art in the measurement of direct, precise stellar masses at pre-main-sequence ages through the analysis of eclipsing binary orbits and circumstellar disk dynamics. We highlight two key issues: (1) The masses determined from disk dynamics require more precise distance determinations that should become available from Gaia soon, and (2) many eclipsing binaries appear disturbed by the presence of tertiary companions that inject heat into and puff up one or both of the inner binary stars, however the dynamical mechanism by which orbital energy is injected as heat remains unknown.
Parabolic orbit determination. Comparison of the Olbers method and algebraic equations
NASA Astrophysics Data System (ADS)
Kuznetsov, V. B.
2016-05-01
In this paper, the Olbers method for the preliminary parabolic orbit determination (in the Lagrange-Subbotin modification) and the method based on systems of algebraic equations for two or three variables proposed by the author are compared. The maximum number of possible solutions is estimated. The problem of selection of the true solution from the set of solutions obtained both using additional equations and by the problem reduction to finding the objective function minimum is considered. The results of orbit determination of the comets 153P/Ikeya-Zhang and 2007 N3 Lulin are cited as examples.
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.
NASA Technical Reports Server (NTRS)
Dunham, J. B.
1980-01-01
An onboard navigation system was developed to aid the design and evaluation of algorithms used in autonomous satellite navigation with Global Positioning System (GPS) data. The performance of the algorithms designed for a GPS Receiver/Processor Assembly (R/PA) intended for LANDSAT-D was investigated during the development phases of the GPS (four to six satellites in the constellation). This evaluation emphasized the effects on the orbit determination accuracy of the expected user clock errors, GPS satellite visibility, force model approximations, and state and covariance propagation approximations. Results are presented giving the sensitivity of orbit determination accuracy to these constraints.
DPOD2005 : Realization of a DORIS terrestrial reference frame for precise orbit determination
NASA Astrophysics Data System (ADS)
Willis, Pascal; Ries, John C.; Soudarin, Laurent; Zelensky, Nikita; Pavlis, Erricos C.
Scientific studies related to altimetry data (mean sea level determination and its time evolution) require centimeter-level orbit determination in the radial component of the satellite. Change in station coordinates and velocities affect the orbit determination and the derived oceanographic results. Following the release of the ITRF2005, we conducted an extensive study related to the DORIS tracking network. For all ground beacons, we verified if the ITRF2005 position and velocity can be extrapolated in time without significant loss of precision. We tried to identified discontinuities in the DORIS coordinates time series, either caused by physical reason, such as Earthquakes, or by instrumental causes. We also identified time periods for which data for a specific station should not be used for orbit determination. In particular, specific stations such as Socorro Island, on which horizontal and vertical movements are detected from the DORIS results will be presented and can be explained by a volcano deformation. Finally, a more complex example will be provided for the Arequipa station, where a major Earthquake happened on June 23, 2001 and for which some relaxation effects are noticeable in the velocity determination even 2 years after the station displacement. A complete set of positions and velocities (by intervals) is given (DPOD2005) and will be used for Jason and TOPEX orbit determination.
NASA Astrophysics Data System (ADS)
Rudenko, Sergei; Gruber, Christian
2016-04-01
This study makes use of current GFZ monthly and daily gravity field products from 2002 to 2014 based on radial basis functions (RBF) instead of time variable gravity field modeling for precise orbit determination of altimetry satellites. Since some monthly solutions are missing in the GFZ GRACE RL05a solution and in order to reach a better quality for the precise orbit determination, daily generated RBF solutions obtained from Kalman filtered GRACE data processing and interpolated in case of gaps have been used. Moreover, since the geopotential coefficients of low degrees are better determined using SLR observations to geodetic satellites like Lageos, Stella, Starlette and Ajisai than from GRACE observations, these terms are co-estimated in the RBF solutions by using apriori SLR-derived values up to degree and order 4. Precise orbits for altimetry satellites Envisat (2002-2012), Jason-1 (2002-2013) and Jason-2 (2008-2014) are then computed over the given time intervals using this approach and compared with the orbits obtained when using other models such as EIGEN-6S4. An analysis of the root-mean-square values of the observation fits of SLR and DORIS observations and the orbit arcs overlaps will allow us to draw a conclusion on the quality of the RBF solution and to use these new trajectories for sea level trend estimates and geophysical application.
Accurate orbit determination strategies for the tracking and data relay satellites
NASA Technical Reports Server (NTRS)
Oza, D. H.; Bolvin, D. T.; Lorah, J. M.; Lee, T.; Doll, C. E.
1995-01-01
The National Aeronautics and Space Administration (NASA) has developed the Tracking and Data Relay Satellite (TDRS) System (TDRSS) for tracking and communications support of low Earth-orbiting satellites. TDRSS has the operational capability of providing 85% coverage for TDRSS-user spacecraft. TDRSS currently consists of five geosynchronous spacecraft and the White Sands Complex (WSC) at White Sands, New Mexico. The Bilateration Ranging Transponder System (BRTS) provides range and Doppler measurements for each TDRS. The ground-based BRTS transponders are tracked as if they were TDRSS-user spacecraft. Since the positions of the BRTS transponders are known, their radiometric tracking measurements can be used to provide a well-determined ephemeris for the TDRS spacecraft. For high-accuracy orbit determination of a TDRSS user, such as the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft, high-accuracy TDRS orbits are required. This paper reports on successive refinements in improved techniques and procedures leading to more accurate TDRS orbit determination strategies using the Goddard Trajectory Determination System (GTDS). These strategies range from the standard operational solution using only the BRTS tracking measurements to a sophisticated iterative process involving several successive simultaneous solutions for multiple TDRSs and a TDRSS-user spacecraft. Results are presented for GTDS-generated TDRS ephemerides produced in simultaneous solutions with the TOPEX/Poseidon spacecraft. Strategies with different user spacecraft, as well as schemes for recovering accurate TDRS orbits following a TDRS maneuver, are also presented. In addition, a comprehensive assessment and evaluation of alternative strategies for TDRS orbit determination, excluding BRTS tracking measurements, are presented.
Precise orbit determination of BeiDou constellation based on BETS and MGEX network.
Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu
2014-01-01
Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 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
NASA Astrophysics Data System (ADS)
Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu
2014-04-01
Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20 cm and 14 cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10 cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved.
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 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.
A Frontier Molecular Orbital determination of the active sites on dispersed metal catalysts
Augustine, R.L.; Lahanas, K.M.
1992-01-01
An angular overlap calculation has been used to determine the s, p and d orbital energy levels of the different types of surface sites present on a dispersed metal catalysts. The basis for these calculations is the reported finding that a large number of catalyzed reactions take place on single atom active sites on the metal surface. Thus, these sites can be considered as surface complexes made up of the central active atom surrounded by near-neighbor metal atom ligands'' with localized surface orbitals perturbed only by these ligands''. These complexes'' are based on a twelve coordinate species with the ligands'' attached to the t{sub 2g} orbitals and the coordinate axes coincident with the direction of the e{sub g} orbitals on the central atom. These data can permit a Frontier Molecular Orbital treatment of specific site activities as long as the surface orbital availability for overlap with adsorbed substrates is considered along with its energy value and symmetry.
A Frontier Molecular Orbital determination of the active sites on dispersed metal catalysts
Augustine, R.L.; Lahanas, K.M.
1992-11-01
An angular overlap calculation has been used to determine the s, p and d orbital energy levels of the different types of surface sites present on a dispersed metal catalysts. The basis for these calculations is the reported finding that a large number of catalyzed reactions take place on single atom active sites on the metal surface. Thus, these sites can be considered as surface complexes made up of the central active atom surrounded by near-neighbor metal atom ``ligands`` with localized surface orbitals perturbed only by these ``ligands``. These ``complexes`` are based on a twelve coordinate species with the ``ligands`` attached to the t{sub 2g} orbitals and the coordinate axes coincident with the direction of the e{sub g} orbitals on the central atom. These data can permit a Frontier Molecular Orbital treatment of specific site activities as long as the surface orbital availability for overlap with adsorbed substrates is considered along with its energy value and symmetry.
New method for on-orbit-determination of parameters for guidance, navigation and control
NASA Astrophysics Data System (ADS)
Clemen, Carsten
2002-07-01
For most scientific satellite missions it is required that the attitude control system works with the greatest possible accuracy to achieve the missions objectives. Therefore it is necessary to know the non-linear system "satellite" with its different parameters e.g. the moments of inertia, the thruster parameters, mass, etc. as exactly as possible. But in most cases these parameters cannot be determined on earth so it has to be done on-orbit. For that reason an applicable estimation method must be found which is robust against all kinds of disturbances acting on the satellite: disturbing torques, noise and offset of the satellites sensors and actuators. Another condition is that only the measurements of the satellite sensors (attitude and angular velocity) can be used for determination. Because of these constraints and the complex non-linear behaviour of the satellite linear and conventional non-linear methods cannot be used for the estimation. A new method must be used. The idea is to compare the sensor measurements which are influenced by the just mentioned disturbances and the results of a computer-aided simulation of the non-disturbed satellite. With the difference between the measured and simulated sensor informations a cost function is determined. In an iteration loop the parameter values are estimated by varying them until this function reaches its minimum. This is performed with the simplex-method by NELDER and MEAD. Here the applicability and accuracy of this method is proved exemplarily by estimating the parameters of a satellites thruster system. This method delivers the parameters sufficiently exact independent from initialized errors in the simulation and disturbances on the real satellite. This method is also applicable to other parameters like for example the moments of inertia or the solar pressure torque, etc.
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.
NASA Astrophysics Data System (ADS)
Guitart, A.; Mesnard, B.
1986-05-01
A satellite tracking campaign was organized, with 4 S-band stations, for 1 wk. The relative geometry of the network with respect to the satellites was an opportunity to show how the most precise orbit can be computed with the operational software. This precise orbit served as a reference to evaluate what can be achieved with one station with range and angular measurements, a typical configuration used for stationkeeping of geostationary satellites. Orbit computation implied numerical integration with gravitational (Earth, Moon, and Sun) and solar radiation pressure forces acting on the satellite. Arc lengths of 2 days gave initial state vectors which were compared every day. Precision of 10 m is achieved. However, an analysis of the influence of parameters in the orbit computations reveals that the absolute accuracy is of the order of 100 m, since modeling perturbations were neglected in the operational software (e.g., polar motion). In a relative sense, the reference orbit allows estimation of systematic errors for other tracking antennas.
Precise orbit determination for GRACE using undifferenced or doubly differenced GPS data
NASA Astrophysics Data System (ADS)
Jäggi, A.; Hugentobler, U.; Bock, H.; Beutler, G.
The two GRACE satellites provide the ideal platform to study the performance of different strategies for precise orbit determination using undifferenced or doubly differenced GPS data. We use pseudo-stochastic orbit modeling techniques in a batch least-squares environment for the two GRACE satellites to outline the mutual benefits of processing doubly differenced instead of undifferenced GPS data. We either process the space baseline only, the space-ground baselines only, or both types of baselines together, and show that the fixing of the GPS double difference carrier phase ambiguities has a significant impact on the space baseline, but also on the space-ground baselines. The validation of the relative orbit positions by inter-satellite K-band observations shows precisions of better than 1 mm in the case of fixed space baseline ambiguities, precisions of a few millimeter in the case of fixed space-ground baseline ambiguities, and precisions of about 1 cm in the case of float ambiguities. We discuss the differences between the various GRACE orbit solutions in order to formulate well suited orbit determination strategies tailored to the GRACE configuration. Satellite laser ranging observations indicate that accuracies between 2 cm and 2.5 cm are achieved.
NASA Technical Reports Server (NTRS)
Daly, J. K.
1974-01-01
The programming techniques used to implement the equations and mathematical techniques of the Houston Operations Predictor/Estimator (HOPE) orbit determination program on the UNIVAC 1108 computer are described. Detailed descriptions are given of the program structure, the internal program structure, the internal program tables and program COMMON, modification and maintainence techniques, and individual subroutine documentation.
Short arc orbit determination for altimeter calibration and validation on TOPEX/POSEIDON
NASA Technical Reports Server (NTRS)
Williams, B. G.; Christensen, E. J.; Yuan, D. N.; Mccoll, K. C.; Sunseri, R. F.
1993-01-01
TOPEX/POSEIDON (T/P) is a joint mission of United States' National Aeronautics and Space Administration (NASA) and French Centre National d'Etudes Spatiales (CNES) design launched August 10, 1992. It carries two radar altimeters which alternately share a common antenna. There are two project designated verification sites, a NASA site off the coast at Pt. Conception, CA and a CNES site near Lampedusa Island in the Mediterranean Sea. Altimeter calibration and validation for T/P is performed over these highly instrumented sites by comparing the spacecraft's altimeter radar range to computed range based on in situ measurements which include the estimated orbit position. This paper presents selected results of orbit determination over each of these sites to support altimeter verification. A short arc orbit determination technique is used to estimate a locally accurate position determination of T/P from less than one revolution of satellite laser ranging (SLR) data. This technique is relatively insensitive to gravitational and non-gravitational force modeling errors and is demonstrated by covariance analysis and by comparison to orbits determined from longer arcs of data and other tracking data types, such as Doppler Orbitography and Radiopositioning Integrated by Satellite (DORIS) and Global Positioning System Demonstration Receiver (GPSDR) data.
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 (Δ DOR) from the CE-2's X-band monitoring and control system experimental are analyzed. It is concluded that the accuracy of Δ 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
NASA Technical Reports Server (NTRS)
Keckler, C. R.; Kibler, K. S.; Powell, L. F.
1979-01-01
A high fidelity simulation of the annular suspension and pointing system (ASPS), its payload, and the shuttle orbiter was used to define the worst case orientations of the ASPS and its payload for the various vehicle disturbances, and to determine the performance capability of the ASPS under these conditions. The most demanding and largest proposed payload, the Solar Optical Telescope was selected for study. It was found that, in all cases, the ASPS more than satisfied the payload's requirements. It is concluded that, to satisfy facility class payload requirements, the ASPS or a shuttle orbiter free-drift mode (control system off) should be utilized.
NASA Technical Reports Server (NTRS)
Vigue, Y.; Lichten, S. M.; Muellerschoen, R. J.; Blewitt, G.; Heflin, M. B.
1993-01-01
Data collected from a worldwide 1992 experiment were processed at JPL to determine precise orbits for the satellites of the Global Positioning System (GPS). A filtering technique was tested to improve modeling of solar-radiation pressure force parameters for GPS satellites. The new approach improves orbit quality for eclipsing satellites by a factor of two, with typical results in the 25- to 50-cm range. The resultant GPS-based estimates for geocentric coordinates of the tracking sites, which include the three DSN sites, are accurate to 2 to 8 cm, roughly equivalent to 3 to 10 nrad of angular measure.
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.
The effects of geopotential resonance on orbit determination for Landsat-4
NASA Technical Reports Server (NTRS)
Hoge, S. L.; Casteel, D. O.; Phenneger, M. C.; Smith, E. A.
1988-01-01
Analysis is presented demonstrating improved performance for Landsat-4 orbit determination using the Goddard Trajectory Determination System with an adjusted Goddard Earth Model-9 (GEM-9) for geopotential coefficients of the 15th degree and order. The orbital state is estimated along with the sine and cosine coefficients of degree and order 15, (C, S) sub 15,15. The estimates are made for two 5-day intervals of range and doppler data, primarily from the Tracking and Data Relay Satellite, during a period of low solar activity in January 1987. The average values of the estimated coefficients (C, S) sub 15,15 are used to modify the GEM-9 model, and orbit determination performance is tested on 17 consecutive 34-hour operational tracking data arcs in January 1987. Significant reductions in the mean values and standard deviations of the along-track position difference and the drag model scaling parameter from solution to solution are observed. The approach is guided by the shallow resonance theory of geopotential orbit perturbations.
Orbit Determination Issues and Results to Incorporate Optical Measurements in Conjunction Operations
NASA Astrophysics Data System (ADS)
Vallado, David A.; Kelso, T. S.; Agapov, Vladimir; Molotov, Igor
2009-03-01
Operations in geosynchronous orbit are important for many aspects of commerce. Avoiding conjunctions between an ever increasingly crowded geosynchronous environment is therefore becoming more important especially in light of the Iridium 33 - Cosmos 2251 collision. SOCRATES has processed Two-Line Element (TLE) set information for over 5 years. Unfortunately, the TLE information is of limited quality, and obtaining high quality ephemerides is difficult. A next step is to see how we can replace the TLE data for those objects for which we do not get operator data (non-participating SOCRATES-GEO oerational satellites or debris). The International Scientific Observing Network (ISON) is an excellent resource to obtain high-quality observations on satellites. The paper introduces the orbit determination, along with test cases and comparisons with known operator orbits. Finally, we discuss how these observations could be used operationally in the conjunction processing and what considerations should be taken into account.
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.
The determination of maximum deep space station slew rates for a high Earth orbiter
NASA Technical Reports Server (NTRS)
Estefan, J. A.
1990-01-01
As developing national and international space ventures, which seek to employ NASA's Deep Space Network (DSN) for tracking and data acquisition, evolve, it is essential for navigation and tracking system analysts to evaluate the operational capability of Deep Space Station antennas. To commission the DSN for use in tracking a highly eccentric Earth orbiter could quite possibly yield the greatest challenges in terms of slewing capability; certainly more so than with a deep-space probe. The focus here is on the determination of the maximum slew rates needed to track a specific high Earth orbiter, namely the Japanese MUSES-B spacecraft of the Very Long Baseline Interferometry Space Observatory Program. The results suggest that DSN 34-m antennas are capable of meeting the slew rate requirements for the nominal MUSES-B orbital geometries currently being considered.
NASA Astrophysics Data System (ADS)
Keating, G. M.; Bougher, S. W.; Theriot, M. E.; Tolson, R. H.; Blanchard, R. C.; Zurek, R. W.; Forbes, J. M.; Murphy, J.
2006-12-01
Designed for aerobraking, Mars Reconnaissance Orbiter (MRO) launched on August 12, 2005, achieved Mars Orbital Insertion (MOI), March 10, 2006, and successfully completed aerobraking on August 30, 2006. Atmospheric density decreases exponentially with increasing height. By small propulsive adjustments of the apoapsis orbital velocity, periapsis altitude was fine tuned to the density surface that safely used the atmosphere of Mars to aerobrake over 445 orbits, providing 890 vertical structures. MRO periapsis precesses from near the South Pole at 6pm LST to near the equator at 3am LST. Meanwhile, apoapsis is brought dramatically from 40,000km at MOI to 480 km at aerobraking completion (ABX). Without aerobraking this would have required an additional 400kg of fuel. After ABX, two small propulsive orbital adjustment maneuvers September 5, 2006 and September 11, 2006 established the final Primary Science Orbit (PSO). Each of the 445 aerobraking orbits provides, a pair of vertical structures inbound toward periapsis and outbound from periapsis, with a distribution of density, scale heights, temperatures, and pressures along the orbital path, providing key in situ insight into various upper atmosphere (> 100 km) processes. One of the major questions for scientists studying Mars is: Where did the water go? Honeywell's substantially improved electronics package for its IMU (QA-2000 accelerometer, gyro, electronics) maximized accelerometer sensitivities at the requests of The George Washington University, JPL, and Lockheed Martin. The improved accelerometer sensitivities allowed density measurements to exceed 200km, at least 40 km higher than with Mars Odyssey (MO). This extends vertical structures from MRO into the neutral lower exosphere, a region where various processes may allow atmospheric gasses to escape. Over the eons, water may have been lost in both the lower atmosphere and the upper atmosphere, thus the water balance throughout the entire atmosphere from
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.
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.
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
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.
TerraSAR-X precise orbit determination with real-time GPS ephemerides
NASA Astrophysics Data System (ADS)
Wermuth, Martin; Hauschild, Andre; Montenbruck, Oliver; Kahle, Ralph
TerraSAR-X is a German Synthetic Aperture Radar (SAR) satellite, which was launched in June 2007 from Baikonour. Its task is to acquire radar images of the Earth's surface. In order to locate the radar data takes precisely, the satellite is equipped with a high-quality dual-frequency GPS receiver -the Integrated Geodetic and Occultation Receiver (IGOR) provided by the GeoForschungsZentrum Potsdam (GFZ). Using GPS observations from the IGOR instrument in a reduced dynamic precise orbit determination (POD), the German Space Operations Center (DLR/GSOC) is computing rapid and science orbit products on a routine basis. The rapid orbit products arrive with a latency of about one hour after data reception with an accuracy of 10-20 cm. Science orbit products are computed with a latency of five days achieving an accuracy of about 5cm (3D-RMS). For active and future Earth observation missions, the availability of near real-time precise orbit information is becoming more and more important. Other applications of near real-time orbit products include the processing of GNSS radio occulation measurements for atmospheric sounding as well as altimeter measurements of ocean surface heights, which are nowadays employed in global weather and ocean circulation models with short latencies. For example after natural disasters it is necessary to evaluate the damage by satellite images as soon as possible. The latency and quality of POD results is mainly driven by the availability of precise GPS ephemerides. In order to have high-quality GPS ephemerides available at real-time, GSOC has developed the real-time clock estimation system RETICLE. The system receives NTRIP-data streams with GNSS observations from the global tracking network of IGS in real-time. Using the known station position, RETICLE estimates precise GPS satellite clock offsets and drifts based on the most recent available IGU predicted orbits. The clock offset estimates have an accuracy of better than 0.3 ns and are
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
Period and Orbital Separation determination of a Subdwarf B Pulsator, EC 20117-4014
NASA Astrophysics Data System (ADS)
Otani, Tomomi; Oswalt, Terry
2016-01-01
EC 20117-4014 (V4640 Sgr) is believed to be a binary system consisting of a pulsating subdwarf B star and a F5V star, however the binary period and orbital distance has not been firmly determined. So far, the most promising theory for the origin of subdwarf B (sdB) stars is that they result from binary mass transfer near the Helium Flush stage. We attempted to constrain this evolutional theory by searching for companions and determining periods and orbital separations around sdB pulsators using the Observed-minus-Calculated (O-C) method. A star's position in space will wobble due to the gravitational forces of any companion. If the star is emitting a periodic signal, its orbital motion around the system's center of mass causes periodic changes in the light pulse arrival times. EC 20117-4014 was monitored from 2010-1 using the 0.6m SARA-CT telescope in Cerro Tololo Inter-American Observatory, Chile. After obtaining the O-C diagrams for the star, useful limits on suspected companions' minimum masses and semimajor axes were calculated. In addition, a modeling experiment was performed to investigate the ranges and combinations of possible companion masses and orbits that are consistent with the observational data. Also, the expected radial velocity semi-amplitude for each O-C companion signal was estimated.
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.
An initial analysis of the data from the Polar Orbiting Geophysical (POGS) Satellite
NASA Technical Reports Server (NTRS)
Langel, R. A.; Sabaka, T. J.; Baldwin, R. T.
1991-01-01
The Polar Orbiting Geophysical Satellite (POGS) was launched in 1990 to measure the geomagnetic field. POGS data from selected magnetically quiet days was chosen, quality checked and deleted where thought to be erroneous. A time and position correction was applied. The resulting data was fit to a degree 13 spherical harmonic model. Evaluation of the quality of the data indicates that it is sufficient for definition of the low degree (approximately less than 8) portion of the geomagnetic field. Further correction of the data time and position may improve this quality.
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.
NASA Astrophysics Data System (ADS)
Murakami, M.
1989-03-01
The subject which this paper deals with is a 1-ppm level determination of the orbits of the Global Positioning System satellites for geodetic applications. A detailed model of the observables is developed. A new method of processing the phase and the range observables simultaneously to determine the GPS orbits is presented. Results are included and discussed.
NASA Technical Reports Server (NTRS)
Goossens, S.; Matsumoto, K.; Noda, H.; Araki, H.; Rowlands, D. D.; Lemoine, F. G.
2011-01-01
The SELENE mission, consisting of three separate satellites that use different terrestrial-based tracking systems, presents a unique opportunity to evaluate the contribution of these tracking systems to orbit determination precision. The tracking data consist of four-way Doppler between the main orbiter and one of the two sub-satellites while the former is over the far side, and of same-beam differential VLBI tracking between the two sub-satellites. Laser altimeter data are also used for orbit determination. The contribution to orbit precision of these different data types is investigated through orbit overlap analysis. It is shown that using four-way and VLBI data improves orbit consistency for all satellites involved by reducing peak values in orbit overlap differences that exist when only standard two-way Doppler and range data are used. Including laser altimeter data improves the orbit precision of the SELENE main satellite further, resulting in very smooth total orbit errors at an average level of 18m. The multi-satellite data have also resulted in improved lunar gravity field models, which are assessed through orbit overlap analysis using Lunar Prospector tracking data. Improvements over a pre-SELENE model are shown to be mostly in the along-track and cross-track directions. Orbit overlap differences are at a level between 13 and 21 m with the SELENE models, depending on whether l-day data overlaps or I-day predictions are used.
Flight dynamics facility operational orbit determination support for the ocean topography experiment
NASA Technical Reports Server (NTRS)
Bolvin, D. T.; Schanzle, A. F.; Samii, M. V.; Doll, C. E.
1991-01-01
The Ocean Topography Experiment (TOPEX/POSEIDON) mission is designed to determine the topography of the Earth's sea surface across a 3 yr period, beginning with launch in June 1992. The Goddard Space Flight Center Dynamics Facility has the capability to operationally receive and process Tracking and Data Relay Satellite System (TDRSS) tracking data. Because these data will be used to support orbit determination (OD) aspects of the TOPEX mission, the Dynamics Facility was designated to perform TOPEX operational OD. The scientific data require stringent OD accuracy in navigating the TOPEX spacecraft. The OD accuracy requirements fall into two categories: (1) on orbit free flight; and (2) maneuver. The maneuver OD accuracy requirements are of two types; premaneuver planning and postmaneuver evaluation. Analysis using the Orbit Determination Error Analysis System (ODEAS) covariance software has shown that, during the first postlaunch mission phase of the TOPEX mission, some postmaneuver evaluation OD accuracy requirements cannot be met. ODEAS results also show that the most difficult requirements to meet are those that determine the change in the components of velocity for postmaneuver evaluation.
Analytical determination of orbital elements using Fourier analysis. I. The radial velocity case
NASA Astrophysics Data System (ADS)
Delisle, J.-B.; Ségransan, D.; Buchschacher, N.; Alesina, F.
2016-05-01
We describe an analytical method for computing the orbital parameters of a planet from the periodogram of a radial velocity signal. The method is very efficient and provides a good approximation of the orbital parameters. The accuracy is mainly limited by the accuracy of the computation of the Fourier decomposition of the signal which is sensitive to sampling and noise. Our method is complementary with more accurate (and more expensive in computer time) numerical algorithms (e.g. Levenberg-Marquardt, Markov chain Monte Carlo, genetic algorithms). Indeed, the analytical approximation can be used as an initial condition to accelerate the convergence of these numerical methods. Our method can be applied iteratively to search for multiple planets in the same system.
NASA Astrophysics Data System (ADS)
Goossens, S. J.; Matsumoto, K.; Kikuchi, F.; Liu, Q.; Hanada, H.; Lemoine, F. G.; Rowlands, D. D.; Ishihara, Y.; Jianguo, Y.; Araki, H.; Noda, H.; Namiki, N.; Iwata, T.
2010-12-01
The Kaguya spacecraft were launched from Tanegashima Space Center on September 14, 2007. Kaguya consists of three orbiters: a main orbiter in a low-altitude (100 km) circular polar orbit, and two sub-satellites (Rstar and Vstar) in elliptical orbits. The satellites were tracked by a variety of terrestrial based tracking systems: in addition to standard two-way Doppler and range tracking, there was 4-way Doppler tracking between Rstar and the main orbiter, providing the first tracking data of a satellite over the lunar far side, and there was same-beam differential VLBI tracking between the two sub-satellites, providing precise orbits for these satellites. The main orbiter was also equipped with a laser altimeter (LALT) to measure the topography of the Moon. At points where the ground tracks of different orbits intersect, these data can provide further constraints on the orbit of the main satellite in the form of crossovers, as essentially the same topography should be measured. This comprehensive data set between the satellites allows for a unique opportunity to evaluate the contribution of these tracking systems to orbit and gravity field determination. Precise orbits are important for geolocation of the topography and camera data, whereas the gravity field can be used for studies of the lunar interior. Here, we present the analysis of the combinations of these tracking data. The use of 4-way and same-beam differential VLBI data leads to large improvements in orbit precision of all satellites involved, where especially peaks in orbit overlap differences during edge-on periods are reduced. The use of the altimetry crossovers improves the orbit of the main satellite further, resulting in an orbit precision of in general less than 20 m. We have also used the full set of SELENE tracking data (including all 4-way and all S-band same-beam differential VLBI data), together with historical data, for gravity field determination. We show a lunar gravity field model with an
Orbit determination and gravitational field accuracy for a Mercury transponder satellite
NASA Technical Reports Server (NTRS)
Vincent, Mark A.; Bender, Pater L.
1990-01-01
Covariance studies were performed to investigate the orbit determination problem for a small transponder satellite in a nearly circular polar orbit with 4-hour period around Mercury. With X band and Ka band Doppler and range measurements, the analysis indicates that the gravitational field through degree and order 10 can be solved for from as few as 40 separate 8-hour arcs of tracking data. In addition, the earth-Mercury distance can be determined during each ranging period with about 6-cm accuracy. The expected geoid accuracy is 10 cm up through degree 5, and 1 m through degree 8. The main error sources were the geocentric range measurement error, the uncertainties in higher degree gravity field terms, which were not solved for, and the solar radiation pressure uncertainty.
Atmospheric drag model for Cassini orbit determination during low altitude Titan flybys
NASA Technical Reports Server (NTRS)
Pelletier, F. J.; Antreasian, P. G.; Bordi, J. J.; Criddle, K. E.; Ionasescu, R.; Jacobson, R. A.; Mackenzie, R. A.; Parcher, D. W.; Stauch, J. R.
2006-01-01
On April 16, 2005, the Cassini spacecraft performed its lowest altitude flyby of Titan to date, the Titan-5 flyby, flying 1027 km above the surface of Titan. This document discusses the development of a Titan atmospheric drag model for the purpose of the orbit determination of Cassini. Results will be presented for the Titan A flyby, the Titan-5 flyby as well as the most recent low altitude Titan flyby, Titan-7. Different solutions will be compared against OD performance in terms of the flyby B-plane parameters, spacecraft thrusting activity and drag estimates. These low altitude Titan flybys were an excellent opportunity to observe the effect of Titan's atmospheric drag on the orbit determination solution and results show that the drag was successfully modeled to provide accurate flyby solutions.
NASA Astrophysics Data System (ADS)
Lei, YANG; Caifa, GUO; Zhengxu, DAI; Xiaoyong, LI; Shaolin, WANG
2016-02-01
The space tracking ship is a moving platform in the TT&C network. The orbit determination precision of the ship plays a key role in the TT&C mission. Based on the measuring data obtained by the ship-borne equipments, the paper presents the mathematic models of the complicated error from the space tracking ship, which can separate the random error and the correction residual error with secondary low frequency from the complicated error. An error simulation algorithm is proposed to analyze the orbit determination precision based on the two set of the different equipments. With this algorithm, a group of complicated error can be simulated from a measured sample. The simulated error groups can meet the requirements of sufficient complicated error for the equipment tests before the mission execution, which is helpful to the practical application.
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.
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
NASA Technical Reports Server (NTRS)
Peters, Palmer N.; Gregory, John C.
1991-01-01
Images produced by pinhole cameras using film sensitive to atomic oxygen provide information on the ratio of spacecraft orbital velocity to the most probable thermal speed of oxygen atoms, provided the spacecraft orientation is maintained stable relative to the orbital direction. Alternatively, as it is described, information on the spacecraft attitude relative to the orbital velocity can be obtained, provided that corrections are properly made for thermal spreading and a co-rotating atmosphere. The LDEF orientation, uncorrected for a co-rotating atmosphere, was determined to be yawed 8.0 minus/plus 0.4 deg from its nominal attitude, with an estimated minus/plus 0.35 deg oscillation in yaw. The integrated effect of inclined orbit and co-rotating atmosphere produces an apparent oscillation in the observed yaw direction, suggesting that the LDEF attitude measurement will indicate even better stability when corrected for a co-rotating atmosphere. The measured thermal spreading is consistent with major exposure occurring during high solar activity, which occurred late during the LDEF mission.
A new method for satellite orbit determination using an operational worldwide transponder network
NASA Technical Reports Server (NTRS)
Lynn, J. J.; Schmid, P. E.; Anderson, R. E.
1974-01-01
The method utilizes computer programs developed for the forthcoming ATS-F/NIMBUS-F tracking and data relay experiment where the basic tracking measurements are multiple path round-trip propagation times and rates. This method of orbit computation has recently been successfully evaluated by tracking a geostationary satellite (ATS-3) using an existing VHF (150 MHz) network of automatic transponders. A master station sequentially interrogates each transponder via the ATS-3. The master site is located at Schenectady, N. Y. and four automatic transponders were located at Shannon, Reykajavik, Buenos Aires, and Seattle respectively. Data at hourly intervals were collected during a 24 hour period on April 18-19, 1973. After correcting this data for known systematic errors it was provided as input to an orbit determination program where all satellite motions during signal propagation are rigorously accounted for. The resulting estimated ATS-3 orbit yielded observational residuals on the order of 100 meters. By using more than one satellite the present scheme is further capable of accurately locating several stationary or mobile terminals as part of the overall orbital solution.
A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver
Wang, Fuhong; Gong, Xuewen; Sang, Jizhang; Zhang, Xiaohong
2015-01-01
Satellite remote sensing systems require accurate, autonomous and real-time orbit determinations (RTOD) for geo-referencing. Onboard Global Positioning System (GPS) has widely been used to undertake such tasks. In this paper, a novel RTOD method achieving decimeter precision using GPS carrier phases, required by China’s HY2A and ZY3 missions, is presented. A key to the algorithm success is the introduction of a new parameter, termed pseudo-ambiguity. This parameter combines the phase ambiguity, the orbit, and clock offset errors of the GPS broadcast ephemeris together to absorb a large part of the combined error. Based on the analysis of the characteristics of the orbit and clock offset errors, the pseudo-ambiguity can be modeled as a random walk, and estimated in an extended Kalman filter. Experiments of processing real data from HY2A and ZY3, simulating onboard operational scenarios of these two missions, are performed using the developed software SATODS. Results have demonstrated that the position and velocity accuracy (3D RMS) of 0.2–0.4 m and 0.2–0.4 mm/s, respectively, are achieved using dual-frequency carrier phases for HY2A, and slightly worse results for ZY3. These results show it is feasible to obtain orbit accuracy at decimeter level of 3–5 dm for position and 0.3–0.5 mm/s for velocity with this RTOD method. PMID:26690149
A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver.
Wang, Fuhong; Gong, Xuewen; Sang, Jizhang; Zhang, Xiaohong
2015-01-01
Satellite remote sensing systems require accurate, autonomous and real-time orbit determinations (RTOD) for geo-referencing. Onboard Global Positioning System (GPS) has widely been used to undertake such tasks. In this paper, a novel RTOD method achieving decimeter precision using GPS carrier phases, required by China's HY2A and ZY3 missions, is presented. A key to the algorithm success is the introduction of a new parameter, termed pseudo-ambiguity. This parameter combines the phase ambiguity, the orbit, and clock offset errors of the GPS broadcast ephemeris together to absorb a large part of the combined error. Based on the analysis of the characteristics of the orbit and clock offset errors, the pseudo-ambiguity can be modeled as a random walk, and estimated in an extended Kalman filter. Experiments of processing real data from HY2A and ZY3, simulating onboard operational scenarios of these two missions, are performed using the developed software SATODS. Results have demonstrated that the position and velocity accuracy (3D RMS) of 0.2-0.4 m and 0.2-0.4 mm/s, respectively, are achieved using dual-frequency carrier phases for HY2A, and slightly worse results for ZY3. These results show it is feasible to obtain orbit accuracy at decimeter level of 3-5 dm for position and 0.3-0.5 mm/s for velocity with this RTOD method. PMID:26690149
Characterizing the three-orbital Hubbard model with determinant quantum Monte Carlo
Kung, Y. F.; Chen, C. -C.; Wang, Yao; Huang, E. W.; Nowadnick, E. A.; Moritz, B.; Scalettar, R. T.; Johnston, S.; Devereaux, T. P.
2016-04-29
Here, we characterize the three-orbital Hubbard model using state-of-the-art determinant quantum Monte Carlo (DQMC) simulations with parameters relevant to the cuprate high-temperature superconductors. The simulations find that doped holes preferentially reside on oxygen orbitals and that the (π,π) antiferromagnetic ordering vector dominates in the vicinity of the undoped system, as known from experiments. The orbitally-resolved spectral functions agree well with photoemission spectroscopy studies and enable identification of orbital content in the bands. A comparison of DQMC results with exact diagonalization and cluster perturbation theory studies elucidates how these different numerical techniques complement one another to produce a more complete understandingmore » of the model and the cuprates. Interestingly, our DQMC simulations predict a charge-transfer gap that is significantly smaller than the direct (optical) gap measured in experiment. Most likely, it corresponds to the indirect gap that has recently been suggested to be on the order of 0.8 eV, and demonstrates the subtlety in identifying charge gaps.« less
Characterizing the three-orbital Hubbard model with determinant quantum Monte Carlo
NASA Astrophysics Data System (ADS)
Kung, Y. F.; Chen, C.-C.; Wang, Yao; Huang, E. W.; Nowadnick, E. A.; Moritz, B.; Scalettar, R. T.; Johnston, S.; Devereaux, T. P.
2016-04-01
We characterize the three-orbital Hubbard model using state-of-the-art determinant quantum Monte Carlo (DQMC) simulations with parameters relevant to the cuprate high-temperature superconductors. The simulations find that doped holes preferentially reside on oxygen orbitals and that the (π ,π ) antiferromagnetic ordering vector dominates in the vicinity of the undoped system, as known from experiments. The orbitally-resolved spectral functions agree well with photoemission spectroscopy studies and enable identification of orbital content in the bands. A comparison of DQMC results with exact diagonalization and cluster perturbation theory studies elucidates how these different numerical techniques complement one another to produce a more complete understanding of the model and the cuprates. Interestingly, our DQMC simulations predict a charge-transfer gap that is significantly smaller than the direct (optical) gap measured in experiment. Most likely, it corresponds to the indirect gap that has recently been suggested to be on the order of 0.8 eV, and demonstrates the subtlety in identifying charge gaps.
Orbit determination across unknown maneuvers using the essential Thrust-Fourier-Coefficients
NASA Astrophysics Data System (ADS)
Ko, Hyun Chul; Scheeres, Daniel J.
2016-01-01
Any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver involving Thrust-Fourier-Coefficients (TFCs). With a selected TFC set as a basis, a thrust acceleration can be constructed to interpolate two unconnected states across an unknown maneuver. This representation technique with TFCs enables us to facilitate the analytical propagation of uncertainties of the satellite state. This approach allows for the usage of existing pre-maneuver orbit estimation to compute the orbit solution after the unknown maneuver. In this paper, we applied this approach to orbit determination (OD) problems across unknown maneuvers by appending different combinations of TFCs to the state vector in the batch filter. The aim is to investigate how different maneuver representations with different TFC sets affect the OD solution across unknown maneuvers. Simulation results show that each TFC set provides different representations of the unknown perturbing acceleration, which yields varying magnitudes of delta velocity for a given maneuver. However, OD solutions across unknown maneuvers using different TFC sets display equivalent performance over the post-maneuver arc as long as those TFC sets are capable of generating the apparent secular motion caused by a given unknown maneuver.
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 Technical Reports Server (NTRS)
Haines, B.; Lichten, S.; Srinivasan, J.; Young, L.
1995-01-01
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.
NASA Technical Reports Server (NTRS)
Goetz, A. F. H.; Rowan, L. C.; Kingston, M. J.
1982-01-01
The Shuttle multispectral IR radiometer (SMIRR) was designed to obtain surface reflectance data in ten spectral bands in order to evaluate the usefulness of a future imaging system for remote mineral identification. Attention was given to the 2.0-2.4 micron region, which has a wealth of spectral absorption features and appeared to have potential for the identification of CO3- and OH-bearing minerals such as the kaolinite and montmorillonite clays. SMIRR radiances were normalized by using a spectrum for dune sand collected in the Kharga Depression in Egypt. Direct identifications have been made of kaolinite-containing and carbonate material, indicating an exceptional potential for future orbital platform narrowband spectral imaging systems for mineralogical mapping.
Orbital Pseudotumor: Uncommon Initial Presentation of IgG4-Related Disease
Carbone, Teresa; Azêdo Montes, Ricardo; Andrade, Beatriz; Lanzieri, Pedro; Mocarzel, Luis
2015-01-01
IgG4-related disease (IgG4-RD) encompasses a group of fibroinflammatory conditions recognized in recent times. The main clinical features include variable degrees of tissue fibrosis, tumorlike expansions, perivascular lymphocytic infiltration rich in IgG4-positive plasma cells, and elevated serum IgG4. A case has been reported of an elderly patient with an unexplained unilateral exophthalmia; biopsy was performed and revealed lymphocytic infiltration, suggesting IgG4-RD. High serum levels of IgG4, in association with a good response to steroid therapy and to the exclusion of other diagnoses, confirmed the hypothesis of orbital pseudotumor by IgG4-RD. PMID:25838962
Determination of the force transmitted by an ion thruster plasma plume to an orbital object
NASA Astrophysics Data System (ADS)
Alpatov, A.; Cichocki, F.; Fokov, A.; Khoroshylov, S.; Merino, M.; Zakrzhevskii, A.
2016-02-01
An approach to determine the force transmitted by the plasma plume of an ion thruster to an orbital object immersed in it using its central projection on a selected plane is proposed. A photo camera is used to obtain the image of the object central projection. The algorithms for the calculation of the transmission of momentum by the impacting ion beam are developed including the determination of the object contour and the correction of the error due to a camera offset from the ion beam axis, and the computation of the fraction of the ion beam that impinges on the object surface.
Precise Orbit Determination of LEO Satellite Using Dual-Frequency GPS Data
NASA Astrophysics Data System (ADS)
Hwang, Yoola; Lee, Byoung-Sun; Kim, Jaehoon; Yoon, Jae-Cheol
2009-06-01
KOrea Multi-purpose SATellite (KOMPSAT)-5 will be launched at 550km altitude in 2010. Accurate satellite position (20 cm) and velocity (0.03 cm/s) are required to treat highly precise Synthetic Aperture Radar (SAR) image processing. Ionosphere delay was eliminated using dual frequency GPS data and double differenced GPS measurement removed common clock errors of both GPS satellites and receiver. SAC-C carrier phase data with 0.1 Hz sampling rate was used to achieve precise orbit determination (POD) with ETRI GNSS Precise Orbit Determination (EGPOD) software, which was developed by ETRI. Dynamic model approach was used and satellite's position, velocity, and the coefficients of solar radiation pressure and drag were adjusted once per arc using Batch Least Square Estimator (BLSE) filter. Empirical accelerations for sinusoidal radial, along-track, and cross track terms were also estimated once per revolution for unmodeled dynamics. Additionally piece-wise constant acceleration for cross-track direction was estimated once per arc. The performance of POD was validated by comparing with JPL's Precise Orbit Ephemeris (POE).
NASA Astrophysics Data System (ADS)
Lubey, D.; Ko, H.; Scheeres, D.
The classical orbit determination (OD) method of dealing with unknown maneuvers is to restart the OD process with post-maneuver observations. However, it is also possible to continue the OD process through such unknown maneuvers by representing those unknown maneuvers with an appropriate event representation. It has been shown in previous work (Ko & Scheeres, JGCD 2014) that any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver connecting those two states using Thrust-Fourier-Coefficients (TFCs). Event representation using TFCs rigorously provides a unique control law that can generate the desired secular behavior for a given unknown maneuver. This paper presents applications of this representation approach to orbit prediction and maneuver detection problem across unknown maneuvers. The TFCs are appended to a sequential filter as an adjoint state to compensate unknown perturbing accelerations and the modified filter estimates the satellite state and thrust coefficients by processing OD across the time of an unknown maneuver. This modified sequential filter with TFCs is capable of fitting tracking data and maintaining an OD solution in the presence of unknown maneuvers. Also, the modified filter is found effective in detecting a sudden change in TFC values which indicates a maneuver. In order to illustrate that the event representation approach with TFCs is robust and sufficiently general to be easily adjustable, different types of measurement data are processed with the filter in a realistic LEO setting. Further, cases with mis-modeling of non-gravitational force are included in our study to verify the versatility and efficiency of our presented algorithm. Simulation results show that the modified sequential filter with TFCs can detect and estimate the orbit and thrust parameters in the presence of unknown maneuvers with or without measurement data during maneuvers. With no measurement
NASA Technical Reports Server (NTRS)
Stephenson, Frank W., Jr.
1988-01-01
The NASA Earth-to-Orbit (ETO) Propulsion Technology Program is dedicated to advancing rocket engine technologies for the development of fully reusable engine systems that will enable space transportation systems to achieve low cost, routine access to space. The program addresses technology advancements in the areas of engine life extension/prediction, performance enhancements, reduced ground operations costs, and in-flight fault tolerant engine operations. The primary objective is to acquire increased knowledge and understanding of rocket engine chemical and physical processes in order to evolve more realistic analytical simulations of engine internal environments, to derive more accurate predictions of steady and unsteady loads, and using improved structural analyses, to more accurately predict component life and performance, and finally to identify and verify more durable advanced design concepts. In addition, efforts were focused on engine diagnostic needs and advances that would allow integrated health monitoring systems to be developed for enhanced maintainability, automated servicing, inspection, and checkout, and ultimately, in-flight fault tolerant engine operations.
NASA Astrophysics Data System (ADS)
Khotkevych, N. V.; Vovk, N. R.; Kolesnichenko, Yu. A.
2016-04-01
A study of electron tunneling from quasi-two-dimensional (surface) states with spin-orbit interaction into bulk-mode states, within the framework of a model of an infinitely thin inhomogeneous tunnel magnetic barrier between two conductors. We analyze how the scattering of quasi-two-dimensional electrons on a single magnetic defect affects the tunneling current in this system. We also obtain an analytical expression for the conductance of the tunnel point-contact, as a function of its distance from the defect. It is shown that analyzing local magnetization oscillations around the defect using spin-polarized scanning tunneling microscopy allows us to determine the spin-orbit interaction constant.
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 Astrophysics Data System (ADS)
Bar-Sever, Yoaz E.; Yunck, Thomas P.; Wu, Sien-Chong
Selective availability (SA) degrades the positioning accuracy for nondifferential users of the GPS Standard Positioning Service (SPS). The often quoted SPS accuracy available under normal conditions is 100 m 2DRMS. In the absence of more specific information, many prospective SPS users adopt the 100 m value in their planning, which exaggerates the error in many cases. SA error is examined for point positioning and dynamic orbit determination for an orbiting user. To minimize SA error, nondifferential users have several options: expand their field of view; observe as many GPS satellites as possible; smooth the error over time; and employ precise GPS ephemerides computed independently, as by NASA and the NGS, rather than the broadcast ephemeris. Simulations show that 3D point position error can be kept to 30 m, and this can be smoothed to 3 m in a few hours.
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 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.
European complement to GPS: Orbit determination strategies and results of preliminary experiments
NASA Astrophysics Data System (ADS)
Deleuze, Muriel; Nouel, François; Escane, Isabelle
1996-06-01
The reference mission of the European Complement to GPS (CE-GPS) is destined to insure the integrity of the navigation message of the U.S. Global Positioning System (GPS) for the Civil Aviation community. CE-GPS is based on a geostationary payload capable of transmitting a GPS-like signal to ground users (civil aviation extended to maritime and terrestrial transport). This signal is broadcasted from a ground station (the Feeder Link Station) towards the geostationary which repeats it. The signal includes both a navigation message (GPS satellites ephemerides) and a health status of the GPS constellation (integrity message). An Operational Orbitography Center capable of processing the orbits of the geostationary and GPS satellites elaborates the navigation and integrity messages. The measurements required for the computation of the orbits are collected by a network of about six Monitoring Stations (in the equatorial area) and transmitted to the Feeder Link station. Experiments are under way to test the different new techniques involved in the development of the ground stations, and to select the optimal strategies for the geostationary orbit determination process and for station synchronization. This paper presents the different orbit determination strategies developed for the geostationary satellite and preliminary results obtained with data collected during two campaigns: autumn 1993 with two ground stations (Paris and Toulouse) and Spring 1994 with three ground stations (Kourou, Hartebeeshoek and Toulouse). The paper is organized in three parts: first, the CE-GPS concept is presented; second, the experimental system designed by CNES to assess the CE-GPS technical performances is described; and last, the experimental results are presented.
NASA Astrophysics Data System (ADS)
Shoji, Mitsuo; Yoshioka, Yasunori; Yamaguchi, Kizashi
2014-07-01
A novel procedure to generate initial broken-symmetry solutions is proposed. Conventional methods for the initial broken-symmetry solutions are the MO alter, HOMO-LUMO mixing and fragment methods. These procedures, however, are quite complex. Our new approach is efficient, automatic and highly practical especially for large QM systems. This approach, called the LNO method, is applied to the following four typical open-shell systems: H2, dicarbene and two iron-sulfur clusters of Rieske-type [2Fe-2S] and [4Fe-4S]. The performance and the efficiency as an electronic structural analysis are discussed. The LNO method will be applicable for general systems in the complicated broken symmetry states.
NASA Technical Reports Server (NTRS)
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.
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 exceess 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 227R.
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.
Modeling of Non-Gravitational Forces for Precise and Accurate Orbit Determination
NASA Astrophysics Data System (ADS)
Hackel, Stefan; Gisinger, Christoph; Steigenberger, Peter; Balss, Ulrich; Montenbruck, Oliver; Eineder, Michael
2014-05-01
Remote sensing satellites support a broad range of scientific and commercial applications. The two radar imaging satellites TerraSAR-X and TanDEM-X provide spaceborne Synthetic Aperture Radar (SAR) and interferometric SAR data with a very high accuracy. The precise reconstruction of the satellite's trajectory is based on the Global Positioning System (GPS) measurements from a geodetic-grade dual-frequency Integrated Geodetic and Occultation Receiver (IGOR) onboard the spacecraft. The increasing demand for precise radar products relies on validation methods, which require precise and accurate orbit products. An analysis of the orbit quality by means of internal and external validation methods on long and short timescales shows systematics, which reflect deficits in the employed force models. Following the proper analysis of this deficits, possible solution strategies are highlighted in the presentation. The employed Reduced Dynamic Orbit Determination (RDOD) approach utilizes models for gravitational and non-gravitational forces. A detailed satellite macro model is introduced to describe the geometry and the optical surface properties of the satellite. Two major non-gravitational forces are the direct and the indirect Solar Radiation Pressure (SRP). The satellite TerraSAR-X flies on a dusk-dawn orbit with an altitude of approximately 510 km above ground. Due to this constellation, the Sun almost constantly illuminates the satellite, which causes strong across-track accelerations on the plane rectangular to the solar rays. The indirect effect of the solar radiation is called Earth Radiation Pressure (ERP). This force depends on the sunlight, which is reflected by the illuminated Earth surface (visible spectra) and the emission of the Earth body in the infrared spectra. Both components of ERP require Earth models to describe the optical properties of the Earth surface. Therefore, the influence of different Earth models on the orbit quality is assessed. The scope of
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.
The challenge of precise orbit determination for STSAT-2C using extremely sparse SLR data
NASA Astrophysics Data System (ADS)
Kim, Young-Rok; Park, Eunseo; Kucharski, Daniel; Lim, Hyung-Chul; Kim, Byoungsoo
2016-03-01
The Science and Technology Satellite (STSAT)-2C is the first Korean satellite equipped with a laser retro-reflector array for satellite laser ranging (SLR). SLR is the only on-board tracking source for precise orbit determination (POD) of STSAT-2C. However, POD for the STSAT-2C is a challenging issue, as the laser measurements of the satellite are extremely sparse, largely due to the inaccurate two-line element (TLE)-based orbit predictions used by the SLR tracking stations. In this study, POD for the STSAT-2C using extremely sparse SLR data is successfully implemented, and new laser-based orbit predictions are obtained. The NASA/GSFC GEODYN II software and seven-day arcs are used for the SLR data processing of two years of normal points from March 2013 to May 2015. To compensate for the extremely sparse laser tracking, the number of estimation parameters are minimized, and only the atmospheric drag coefficients are estimated with various intervals. The POD results show that the weighted root mean square (RMS) post-fit residuals are less than 10 m, and the 3D day boundaries vary from 30 m to 3 km. The average four-day orbit overlaps are less than 20/330/20 m for the radial/along-track/cross-track components. The quality of the new laser-based prediction is verified by SLR observations, and the SLR residuals show better results than those of previous TLE-based predictions. This study demonstrates that POD for the STSAT-2C can be successfully achieved against extreme sparseness of SLR data, and the results can deliver more accurate predictions.
Orbit Determination of Chang'e-3 and Positioning of the Lander and the Rover
NASA Astrophysics Data System (ADS)
Huang, Y.; Chang, S.; Li, P.; Hu, X.
2014-12-01
The Chang'E-3 (CE-3) lunar probe of China was launched on 2 December 2013. After about 112 h of flight, it was captured by the Moon on 6 December, and entered a polar, near circular lunar orbit with an altitude of approximately 100 km. The probe's flight on 100 km*100 km and 100 km*15 km orbit lasted about 4 days respectively, then the probe soft landed on the east of Sinus Iridum area at 13:11 UTC on 14 December successfully. Results on precision orbit determination and positioning of the lander and the rover are presented here. We describe the data, modeling and methods used to achieve position knowledge. In addition to the radiometric X-band range and Doppler tracking data, Delta Differential One-way Ranging (ΔDOR) data are also used in the calculation, which shows that they can improve the accuracy of the orbit reconstruction. Total position overlap differences are about 20 m and 30 m for the 100 km*100 km and 100 km*15 km lunar orbit respectively, increased by ~50 % with respect to CE-2. A kinematic statistical method is applied to determine the position of the lander and relative position of the rover with respect to the lander. The location of the lander is computed as: 44.1216º N, 19.5124º W and -2632.0 m in the lunar Mean Axes coordinate system. The position difference of the lander is better than 50 m compared to the result of the LRO photograph. From 15 to 21 December, the rover walked around the lander, and took photos of each other at the parking point A, B, C, D, E (max distance from the lander is about 25 m). The delta VLBI phase delay data are used to compute the relative position of the rover at the parking points, and the accuracy of the relative position can reach to 1-2 m comparing with the results of visual method.
Orbit of the OJ287 black hole binary as determined from the General Relativity centenary flare
NASA Astrophysics Data System (ADS)
Valtonen, Mauri; Gopakumar, Achamveedu; Mikkola, Seppo; Zola, Staszek; Ciprini, Stefano; Matsumoto, Katsura; Sadakane, Kozo; Kidger, Mark; Gazeas, Kosmas; Nilsson, Kari; Berdyugin, Andrei; Piirola, Vilppu; Jermak, Helen; Baliyan, Kiran; Hudec, Rene; Reichart, Daniel
2016-05-01
OJ287 goes through large optical flares twice each 12 years. The times of these flares have been predicted successfully now 5 times using a black hole binary model. In this model a secondary black hole goes around a primary black hole, impacting the accretion disk of the latter twice per orbital period, creating a thermal flare. Together with 6 flares from the historical data base, the set of flare timings determines uniquely the 7 parameters of the model: the two masses, the primary spin, the major axis, eccentricity and the phase of the orbit, plus a time delay parameter that gives the extent of time between accretion disk impacts and the related optical flares. Based on observations by the OJ287-15/16 Collaboration, OJ287 went into the phase of rapid flux rise on November 25, on the centenary of Einstein’s General Relativity, and peaked on December 5. At that time OJ287 was the brightest in over 30 years in optical wavelengths. The flare was of low polarization, and did not extend beyond the optical/UV region of the spectrum. On top of the main flare there were a number of small flares; their excess brightness correlates well with the simultaneous X-ray data. With these properties the main flare qualifies as the marker of the orbit of the secondary going around the primary black hole. Since the orbit solution is strongly over-determined, its parameters are known very accurately, at better than one percent level for the masses and the spin. The next flare is predicted to peak on July 28, 2019.Detailed monitoring of this event should allow us to test, for the first time, the celebrated black hole no-hair theorem for a massive black hole at the 10% level. The present data is consistent with the theorem only at a 30% level. The main difficulty in observing OJ287 from Earth at our predicted epoch is its closeness to the sun. Therefore, it is desirable to monitor OJ287 from a space-based telescope not in the vicinity of Earth. Unfortunately, this unique opportunity
Precision Time Transfer and Obit Determination Using Laser Ranging to Lunar Reconnaissance Orbiter
NASA Astrophysics Data System (ADS)
Mao, D.; Barker, M. K.; Clarke, C. B.; Golder, J. E.; Hoffman, E.; Horvath, J. E.; Mazarico, E.; Mcgarry, J.; Neumann, G. A.; Torrence, M. H.; Rowlands, D. D.; Skillman, D.; Smith, D. E.; Sun, X.; Zuber, M. T.
2011-12-01
Since the commissioning of LRO in June, 2009, one-way laser ranging (LR) to Lunar Reconnaissance Orbiter (LRO) has been conducted successfully from NASA's Next Generation Satellite Laser Ranging System (NGSLR) at Goddard Geophysical and Astronomical observatory (GGAO) in Greenbelt, Maryland. With the support of the International Laser Ranging Service (ILRS), ten international satellite laser ranging (SLR) ground stations have participated in this experiment and over 1200 hours of ranging data have been collected. In addition to supplementing the precision orbit determination (POD) of LRO, LR is able to perform time transfer between the ground station and the spacecraft clocks. The LRO clock oscillator is stable to 1 part in 10^{12} over several hours, and as stable for much longer periods after correcting for a long-term drift rate and an aging rate. With a precisely-determined LRO ephemeris, the oscillator-determined laser pulse receive time can be differenced with ground station clock transmit times using H-maser and GPS-steered Rb oscillators as references. Simultaneous ranging to LRO among 2, 3, or 4 ground stations has made it possible for relative time transfer among the participating LR stations. Results have shown about 100 ns difference between some LR stations and the primary NGSLR station. At present, the time transfer accuracy is limited to 100 ns at NGSLR. However, an All-View GPS receiver has been installed, which, in combination with a H-maser, is expected to improve the accuracy to 1 ns r.m.s. at NGSLR. Results of new ranging and time transfer experiments using the new time base will be reported. The ability to use LR for time transfer validates the selection of a commercially-supplied, oven-controlled crystal oscillator on board LRO for one-way laser ranging.The increased clock accuracy also provides stronger orbit constraints for LRO POD. The improvements due to including LR data in the LRO POD will be presented.
NASA Technical Reports Server (NTRS)
Rourke, K. H.; Jordan, J. F.
1972-01-01
This paper presents the results of the applications of advanced filtering methods to the determination of the interplanetary orbit of the Mariner '71 spacecraft. The advanced techniques are specific extensions of the Kalman filter. The special problems associated with applying these techniques are discussed and the particular algorithmic implementations are outlined. The advanced methods are compared against the weighted least squares filters of conventional application. The results reveal that relatively simple advanced filter configurations yield solutions superior to those of the conventional methods when applied to the Mariner '71 radio measurements.
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.
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
Code of Federal Regulations, 2010 CFR
2010-07-01
..., the amortization of initial liabilities, and the allocation fraction. 4211.36 Section 4211.36 Labor... initial liabilities, and the allocation fraction. (a) General rule. A plan using any of the allocation... participation under their prior plans. An amendment under this paragraph must include an allocation...
Code of Federal Regulations, 2011 CFR
2011-07-01
..., the amortization of initial liabilities, and the allocation fraction. 4211.36 Section 4211.36 Labor... initial liabilities, and the allocation fraction. (a) General rule. A plan using any of the allocation... participation under their prior plans. An amendment under this paragraph must include an allocation...
NASA Astrophysics Data System (ADS)
Centinello, F. J.; Zuber, M. T.; Mazarico, E.
2013-12-01
The Dawn spacecraft orbited the protoplanet Vesta from May 3, 2011 to July 25, 2012. Precise orbit determination was critical for the geophysical investigation, as well as the definition of the Vesta-fixed reference frame and the subsequent registration of datasets to the surface. GEODYN, the orbit determination and geodetic parameter estimation software of NASA Goddard Spaceflight Center, was used to compute the orbit of the Dawn spacecraft and estimate the gravity field of Vesta. GEODYN utilizes radiometric Doppler and range measurements, and was modified to process image data from Dawn's cameras. X-band radiometric measurements were acquired by the NASA Deep Space Network (DSN). The addition of the capability to process image constraints decreases position uncertainty in the along- and cross-orbit track directions because of their geometric strengths compared with radiometric measurements. This capability becomes critical for planetary missions such as Dawn due to the weak gravity environment, where non-conservative forces affect the orbit more than typical of orbits at larger planetary bodies. Radiometric measurements were fit to less than 0.1 mm/s and 5 m for Doppler and range during the Survey orbit phase (compared with measurement noise RMS of about 0.05 mm/s and 2 m for Doppler and range). Image constraint RMS was fit to less than 100 m (resolution is 5 - 150 m/pixel, depending on the spacecraft altitude). Orbits computed using GEODYN were used to estimate a 20th degree and order gravity field of Vesta. The quality of the orbit determination and estimated gravity field with and without image constraints was assessed through comparison with the spacecraft trajectory and gravity model provided by the Dawn Science Team.
20 CFR 405.201 - Reviewing an initial determination-general.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false Reviewing an initial determination-general... PROCESS FOR ADJUDICATING INITIAL DISABILITY CLAIMS Review of Initial Determinations by a Federal Reviewing Official § 405.201 Reviewing an initial determination—general. If you are dissatisfied with the...
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
Phase center modeling for LEO GPS receiver antennas and its impact on precise orbit determination
NASA Astrophysics Data System (ADS)
Jäggi, Adrian; Dach, R.; Montenbruck, O.; Hugentobler, U.; Bock, H.; Beutler, G.
2009-12-01
Most satellites in a low-Earth orbit (LEO) with demanding requirements on precise orbit determination (POD) are equipped with on-board receivers to collect the observations from Global Navigation Satellite systems (GNSS), such as the Global Positioning System (GPS). Limiting factors for LEO POD are nowadays mainly encountered with the modeling of the carrier phase observations, where a precise knowledge of the phase center location of the GNSS antennas is a prerequisite for high-precision orbit analyses. Since 5 November 2006 (GPS week 1400), absolute instead of relative values for the phase center location of GNSS receiver and transmitter antennas are adopted in the processing standards of the International GNSS Service (IGS). The absolute phase center modeling is based on robot calibrations for a number of terrestrial receiver antennas, whereas compatible antenna models were subsequently derived for the remaining terrestrial receiver antennas by conversion (from relative corrections), and for the GNSS transmitter antennas by estimation. However, consistent receiver antenna models for space missions such as GRACE and TerraSAR-X, which are equipped with non-geodetic receiver antennas, are only available since a short time from robot calibrations. We use GPS data of the aforementioned LEOs of the year 2007 together with the absolute antenna modeling to assess the presently achieved accuracy from state-of-the-art reduced-dynamic LEO POD strategies for absolute and relative navigation. Near-field multipath and cross-talk with active GPS occultation antennas turn out to be important and significant sources for systematic carrier phase measurement errors that are encountered in the actual spacecraft environments. We assess different methodologies for the in-flight determination of empirical phase pattern corrections for LEO receiver antennas and discuss their impact on POD. By means of independent K-band measurements, we show that zero-difference GRACE orbits can be
Ab Initio determination of Cu 3d orbital energies in layered copper oxides
Hozoi, Liviu; Siurakshina, Liudmila; Fulde, Peter; van den Brink, Jeroen
2011-01-01
It has long been argued that the minimal model to describe the low-energy physics of the high Tc superconducting cuprates must include copper states of other symmetries besides the canonical one, in particular the orbital. Experimental and theoretical estimates of the energy splitting of these states vary widely. With a novel ab initio quantum chemical computational scheme we determine these energies for a range of copper-oxides and -oxychlorides, determine trends with the apical Cu–ligand distances and find excellent agreement with recent Resonant Inelastic X-ray Scattering measurements, available for La2CuO4, Sr2CuO2Cl2, and CaCuO2. PMID:22355584
NASA Astrophysics Data System (ADS)
Fang, Haijian; Zhang, Rongzhi; Wang, Jiasong; Wang, Dan; Guo, Hai
2015-10-01
The injected transfer orbit of lunar probe Chang'E 5T1 (CE-5T1) is determined immediately after the probe separates from its launcher. As the first orbit in the lunar flight, the CE-5T1 injected transfer orbit is crucial to the consequence of rocket vehicle launch mission and the probe's subsequent midway orbital manoeuvre. In this paper, we discuss the problem of using rocket GPS measurements to determine the probe velocity increment due to mechanical separation, and subsequently the injected transfer orbit determination of CE-5T1. Motivated by the post-mission analysis of lunar probe Chang'E 3 (CE-3), we give theoretical evidence to explain the physical phenomenon of semi-major axis sudden change at the probe separation instant through the derivation of the Vis-Viva equation. In succession, we focus on the description of the procedure used for the orbit determination performed on separated arcs of rocket GPS measurements through the use of momentum conservation to determine the probe separation velocity. Finally, actual flight data of the CE-3 and CE-5T1 missions are used for the validation.
NASA Astrophysics Data System (ADS)
Peng, Dong-ju; Wu, Bin
2012-10-01
With the precise GPS ephemeris and clock error available, the iono- spheric delay is left as the dominant error source in the single-frequency GPS data. Thus, the removal of ionospheric effects is a ma jor prerequisite for an improved orbit reconstruction of LEO satellites based on the single-frequency GPS data. In this paper, the use of Global Ionospheric Maps (GIM) in kine- matic and dynamic orbit determinations for LEO satellites with single-frequency GPS pseudorange measurements is discussed first, and then, estimating the iono- spheric scale factor to remove the ionospheric effects from the C/A-code pseu- dorange measurements for both kinematic and dynamic orbit determinations is addressed. As it is known that the ionospheric delay of space-borne GPS sig- nals is strongly dependent on the orbit altitudes of LEO satellites, we select the real C/A-code pseudorange measurement data of the CHAMP, GRACE, TerraSAR-X and SAC-C satellites with altitudes between 300 km and 800 km as sample data in this paper. It is demonstrated that the approach to eliminating ionospheric effects in C/A-code pseudorange measurements by estimating the ionospheric scale factor is highly effective. Employing this approach, the accu- racy of both kinematic and dynamic orbits can be improved notably. Among those five LEO satellites, CHAMP with the lowest orbit altitude has the most remarkable improvements in orbit accuracy, which are 55.6% and 47.6% for kine- matic and dynamic orbits, respectively. SAC-C with the highest orbit altitude has the least improvements in orbit accuracy accordingly, which are 47.8% and 38.2%, respectively.
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2010 CFR
2010-10-01
... (HIPAA) and related CMS manual 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... national use by covered entities under HIPAA; and (vi) Any other requirements specified by CMS....
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2012 CFR
2012-10-01
... (HIPAA) and related CMS manual 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... national use by covered entities under HIPAA; and (vi) Any other requirements specified by CMS....
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2011 CFR
2011-10-01
... (HIPAA) and related CMS manual 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... national use by covered entities under HIPAA; and (vi) Any other requirements specified by CMS....
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2014 CFR
2014-10-01
... (HIPAA) and related CMS manual 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... national use by covered entities under HIPAA; and (vi) Any other requirements specified by CMS....
42 CFR 405.921 - Notice of initial determination.
Code of Federal Regulations, 2013 CFR
2013-10-01
... (HIPAA) and related CMS manual 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... national use by covered entities under HIPAA; and (vi) Any other requirements specified by CMS....
NASA Astrophysics Data System (ADS)
Desmars, J.; Camargo, J. I. B.; Braga-Ribas, F.; Vieira-Martins, R.; Assafin, M.; Vachier, F.; Colas, F.; Ortiz, J. L.; Duffard, R.; Morales, N.; Sicardy, B.; Gomes-Júnior, A. R.; Benedetti-Rossi, G.
2015-12-01
Context. The prediction of stellar occultations by trans-Neptunian objects (TNOs) and Centaurs is a difficult challenge that requires accuracy both in the occulted star position and in the object ephemeris. Until now, the most used method of prediction, involving dozens of TNOs/Centaurs, has been to consider a constant offset for the right ascension and for the declination with respect to a reference ephemeris, usually the latest public version. This offset is determined as the difference between the most recent observations of the TNO/Centaur and the reference ephemeris. This method can be successfully applied when the offset remains constant with time, i.e. when the orbit is stable enough. In this case, the prediction even holds for occultations that occur several days after the last observations. Aims: This paper presents an alternative method of prediction, based on a new accurate orbit determination procedure, which uses all the available positions of the TNO from the Minor Planet Center database, as well as sets of new astrometric positions from unpublished observations. Methods: Orbits were determined through a numerical integration procedure called NIMA, in which we developed a specific weighting scheme that considers the individual precision of the observation, the number of observations performed during one night by the same observatory, and the presence of systematic errors in the positions. Results: The NIMA method was applied to 51 selected TNOs and Centaurs. For this purpose, we performed about 2900 new observations in several observatories (European South Observatory, Observatório Pico dos Dias, Pic du Midi, etc.) during the 2007-2014 period. Using NIMA, we succeed in predicting the stellar occultations of 10 TNOs and 3 Centaurs between July 2013 and February 2015. By comparing the NIMA and Jet Propulsion Laboratory (JPL) ephemerides, we highlight the variation in the offset between them with time, by showing that, generally, the constant offset
Orbit correction in an orbit separated cyclotron
NASA Astrophysics Data System (ADS)
Plostinar, C.; Rees, G. H.
2014-04-01
The orbit separated proton cyclotron (OSC) described in [1] differs in concept from that of a separated orbit cyclotron (SOC) [2]. Synchronous acceleration in an OSC is based on harmonic number jumps and orbit length adjustments via reverse bending. Four-turn acceleration in the OSC enables it to have four times fewer cryogenic-cavity systems than in a superconducting linac of the same high beam power and energy range. Initial OSC studies identified a progressive distortion of the spiral beam orbits by the off-axis, transverse deflecting fields in its accelerating cavities. Compensation of the effects of these fields involves the repeated use of a cavity field map, in a 3-D linac tracking code, to determine the modified arc bends required for the OSC ring. Subsequent tracking studies confirm the compensation scheme and show low emittance growth in acceleration.
Frozen Orbital Plane Solutions for Satellites in Nearly Circular Orbit
NASA Astrophysics Data System (ADS)
Ulivieri, Carlo; Circi, Christian; Ortore, Emiliano; Bunkheila, Federico; Todino, Francesco
2013-08-01
This paper deals with the determination of the initial conditions (right ascension of the ascending node and inclination) that minimize the orbital plane variation for nearly circular orbits with a semimajor axis between 3 and 10 Earth radii. An analysis of two-line elements over the last 40 years for mid-, geostationary-, and high-Earth orbits has shown, for initially quasi-circular orbits, low eccentricity variations up to the geostationary altitude. This result makes the application of mathematical models based on satellite circular orbits advantageous for a fast prediction of long-term temporal evolution of the orbital plane. To this purpose, a previous model considering the combined effect due to the Earth's oblateness, moon, and sun (both in circular orbit) has been improved in terms of required computational time and accuracy. The eccentricity of the sun and moon and the equinoctial precession have been taken into account. Resonance phenomena with the lunar plane motion have been found in mid-Earth orbit. Dynamical properties concerning the precession motions of the orbital pole have been investigated, and frozen solutions for geosynchronous and navigation satellites have been proposed. Finally, an accurate model validation has also been carried out by comparing the obtained results with two-line elements of abandoned geostationary-Earth orbit and mid-Earth orbit satellites.
GPS-Based Precision Orbit Determination for a New Era of Altimeter Satellites: Jason-1 and ICESat
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.; Rowlands, David D.; Lemoine, Frank G.; Zelensky, Nikita P.; Williams, Teresa A.
2003-01-01
Accurate positioning of the satellite center of mass is necessary in meeting an altimeter mission's science goals. The fundamental science observation is an altimetric derived topographic height. Errors in positioning the satellite's center of mass directly impact this fundamental observation. Therefore, orbit error is a critical Component in the error budget of altimeter satellites. With the launch of the Jason-1 radar altimeter (Dec. 2001) and the ICESat laser altimeter (Jan. 2003) a new era of satellite altimetry has begun. Both missions pose several challenges for precision orbit determination (POD). The Jason-1 radial orbit accuracy goal is 1 cm, while ICESat (600 km) at a much lower altitude than Jason-1 (1300 km), has a radial orbit accuracy requirement of less than 5 cm. Fortunately, Jason-1 and ICESat POD can rely on near continuous tracking data from the dual frequency codeless BlackJack GPS receiver and Satellite Laser Ranging. Analysis of current GPS-based solution performance indicates the l-cm radial orbit accuracy goal is being met for Jason-1, while radial orbit accuracy for ICESat is well below the 54x1 mission requirement. A brief overview of the GPS precision orbit determination methodology and results for both Jason-1 and ICESat are presented.
NASA Astrophysics Data System (ADS)
Hassler, D. M.; Deforest, C.; Spice Team
2011-12-01
At large spatial scales, the structure of the solar wind and it's mapping back to the solar corona, is thought to be reasonably well understood. However, the detailed structure of the various source regions at chromospheric and transition region heights is extremely complex, and less well understood. Determining this connection between heliospheric structures and their source regions at the Sun is one of the overarching objective of the Solar Orbiter mission. During perihelion segments of its orbit, when the spacecraft is in quasi-corotation with the Sun, Solar Orbiter will determine the plasma parameters and compositional signatures of the solar wind, which can be compared directly with the spectroscopic signatures of coronal ions with differing charge-to-mass ratios and FIP. One of the key instruments on the Solar Orbiter mission to make these remote sensing measurements is the SPICE (Spectral Imaging of the Coronal Environment) imaging spectrograph. SPICE will provide the images and plasma diagnostics needed to characterize the plasma state in different source regions, from active regions to quiet Sun to coronal holes. By comparing composition, plasma parameters, and low/high FIP ratios of structures remotely, with those measured directly at the Solar Orbiter spacecraft, Solar Orbiter will provide the first direct link between solar wind structures and their source regions at the Sun. This talk will provide a background of previous compositional correlation measurements and an outline of the method to be used for comparing the spectroscopic and in-situ plasma parameters to be measured with Solar Orbiter.
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.
Real-Time Orbit Determination for Future Korean Regional Navigation Satellite System
NASA Astrophysics Data System (ADS)
Shin, Kihae; Oh, Hyungjik; Park, Sang-Young; Park, Chandeok
2016-03-01
This paper presents an algorithm for Real-Time Orbit Determination (RTOD) of navigation satellites for the Korean Regional Navigation Satellite System (KRNSS), when the navigation satellites generate ephemeris by themselves in abnormal situations. The KRNSS is an independent Regional Navigation Satellite System (RNSS) that is currently within the basic/preliminary research phase, which is intended to provide a satellite navigation service for South Korea and neighboring countries. Its candidate constellation comprises three geostationary and four elliptical inclined geosynchronous orbit satellites. Relative distance ranging between the KRNSS satellites based on Inter-Satellite Ranging (ISR) is adopted as the observation model. The extended Kalman filter is used for real-time estimation, which includes fine-tuning the covariance, measurement noise, and process noise matrices. Simulation results show that ISR precision of 0.3-0.7 m, ranging capability of 65,000 km, and observation intervals of less than 20 min are required to accomplish RTOD accuracy to within 1 m. Furthermore, close correlation is confirmed between the dilution of precision and RTOD accuracy.
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.
NASA Astrophysics Data System (ADS)
Guo, Jing; Xu, Xiaolong; Zhao, Qile; Liu, Jingnan
2016-02-01
This contribution summarizes the strategy used by Wuhan University (WHU) to determine precise orbit and clock products for Multi-GNSS Experiment (MGEX) of the International GNSS Service (IGS). In particular, the satellite attitude, phase center corrections, solar radiation pressure model developed and used for BDS satellites are addressed. In addition, this contribution analyzes the orbit and clock quality of the quad-constellation products from MGEX Analysis Centers (ACs) for a common time period of 1 year (2014). With IGS final GPS and GLONASS products as the reference, Multi-GNSS products of WHU (indicated by WUM) show the best agreement among these products from all MGEX ACs in both accuracy and stability. 3D Day Boundary Discontinuities (DBDs) range from 8 to 27 cm for Galileo-IOV satellites among all ACs' products, whereas WUM ones are the largest (about 26.2 cm). Among three types of BDS satellites, MEOs show the smallest DBDs from 10 to 27 cm, whereas the DBDs for all ACs products are at decimeter to meter level for GEOs and one to three decimeter for IGSOs, respectively. As to the satellite laser ranging (SLR) validation for Galileo-IOV satellites, the accuracy evaluated by SLR residuals is at the one decimeter level with the well-known systematic bias of about -5 cm for all ACs. For BDS satellites, the accuracy could reach decimeter level, one decimeter level, and centimeter level for GEOs, IGSOs, and MEOs, respectively. However, there is a noticeable bias in GEO SLR residuals. In addition, systematic errors dependent on orbit angle related to mismodeled solar radiation pressure (SRP) are present for BDS GEOs and IGSOs. The results of Multi-GNSS combined kinematic PPP demonstrate that the best accuracy of position and fastest convergence speed have been achieved using WUM products, particularly in the Up direction. Furthermore, the accuracy of static BDS only PPP degrades when the BDS IGSO and MEO satellites switches to orbit-normal orientation
NASA Astrophysics Data System (ADS)
Son, Ju Young; Jo, Jung Hyun; Choi, Jin
2015-09-01
To protect and manage the Korean space assets including satellites, it is important to have precise positions and orbit information of each space objects. While Korea currently lacks optical observatories dedicated to satellite tracking, the Korea Astronomy and Space Science Institute (KASI) is planning to establish an optical observatory for the active generation of space information. However, due to geopolitical reasons, it is difficult to acquire an adequately sufficient number of optical satellite observatories in Korea. Against this backdrop, this study examined the possible locations for such observatories, and performed simulations to determine the differences in precision of optical orbit estimation results in relation to the relative baseline distance between observatories. To simulate more realistic conditions of optical observation, white noise was introduced to generate observation data, which was then used to investigate the effects of baseline distance between optical observatories and the simulated white noise. We generated the optical observations with white noise to simulate the actual observation, estimated the orbits with several combinations of observation data from the observatories of various baseline differences, and compared the estimated orbits to check the improvement of precision. As a result, the effect of the baseline distance in combined optical GEO satellite observation is obvious but small compared to the observation resolution limit of optical GEO observation.
NASA Astrophysics Data System (ADS)
Lala, P.
1981-04-01
Papers are presented on the use of point mass models of the geopotential for orbit predictions, on earth ocean tides from long-term analysis of satellite orbits, on the motion of an artificial satellite under the terrestrial radiation pressure, and on the generation of satellite position (and velocity) by a mixed analytical-numerical procedure. Attention is also given to the possibilities of determining the influence of earth body tides on the motion of artificial satellites and to a general time element for orbit integration in Cartesian coordinates.
Determinants of breastfeeding initiation among mothers in Kuwait
2010-01-01
Background Exclusive breastfeeding is recommended as the optimal way to feed infants for the first six months of life. While overall breastfeeding rates are high, exclusive breastfeeding is relatively uncommon among Middle Eastern women. The objective of this study was to identify the incidence of breastfeeding amongst women in the six governorates of Kuwait and the factors associated with the initiation of breastfeeding. Methods A sample of 373 women (aged 17-47 years), recruited shortly after delivery from four hospitals in Kuwait, completed a structured, interviewer-administered questionnaire. Multivariate logistic regression analysis was used to identify those factors independently associated with the initiation of breastfeeding. Results In total, 92.5% of mothers initiated breastfeeding and at discharge from hospital the majority of mothers were partially breastfeeding (55%), with only 30% of mothers fully breastfeeding. Prelacteal feeding was the norm (81.8%) and less than 1 in 5 infants (18.2%) received colostrum as their first feed. Only 10.5% of infants had been exclusively breastfed since birth, the remainder of the breastfed infants having received either prelacteal or supplementary infant formula feeds at some time during their hospital stay. Of the mothers who attempted to breastfeed, the majority of women (55.4%) delayed their first attempt to breastfeed until 24 hours or more after delivery. Breastfeeding at discharge from hospital was positively associated with paternal support for breastfeeding and negatively associated with delivery by caesarean section and with the infant having spent time in the Special Care Nursery. Conclusions The reasons for the high use of prelacteal and supplementary formula feeding warrant investigation. Hospital policies and staff training are needed to promote the early initiation of breastfeeding and to discourage the unnecessary use of infant formula in hospital, in order to support the establishment of exclusive
Validity of repeated initial rise thermoluminescence kinetic parameter determinations
Kierstead, J.A.; Levy, P.W.
1990-01-01
The validity of thermoluminescence (TL) analysis by repeated initial rise measurements has been studied by computer simulation. Thermoluminescence described by 1st Order, 2nd Order, General One Trap and Interactive TL Kinetics was investigated. In the simulation each of the repeated temperature increase and decrease cycles contains a linear temperature increase followed by a decrease appropriate for radiative cooling, i.e. the latter is approximated by a decreasing exponential. The activation energies computed from the simulated emission are readily compared with those used to compute the TL emission. In all cases studied, the repeated initial rise technique provides reliable results only for single peak glow curves or for glow curves containing peaks that do not overlap and, if sufficiently separated, the lowest temperature peak in multipeak curves. Also the temperatures, or temperature cycles corresponding to correct activation energies occur on the low temperature side of the normal glow curve, often well below the peak temperature. A variety of misleading and/or incorrect results an be obtained when the repeated initial rise technique is applied to TL systems that produce overlapping peaks in the usual glow curve. 6 refs., 10 figs.
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.
A numerical comparison of discrete Kalman filtering algorithms - An orbit determination case study
NASA Technical Reports Server (NTRS)
Thornton, C. L.; Bierman, G. J.
1976-01-01
An improved Kalman filter algorithm based on a modified Givens matrix triangularization technique is proposed for solving a nonstationary discrete-time linear filtering problem. The proposed U-D covariance factorization filter uses orthogonal transformation technique; measurement and time updating of the U-D factors involve separate application of Gentleman's fast square-root-free Givens rotations. Numerical stability and accuracy of the algorithm are compared with those of the conventional and stabilized Kalman filters and the Potter-Schmidt square-root filter, by applying these techniques to a realistic planetary navigation problem (orbit determination for the Saturn approach phase of the Mariner Jupiter-Saturn Mission, 1977). The new algorithm is shown to combine the numerical precision of square root filtering with the efficiency of the original Kalman algorithm.
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.
A numerical comparison of discrete Kalman filtering algorithms: An orbit determination case study
NASA Technical Reports Server (NTRS)
Thornton, C. L.; Bierman, G. J.
1976-01-01
The numerical stability and accuracy of various Kalman filter algorithms are thoroughly studied. Numerical results and conclusions are based on a realistic planetary approach orbit determination study. The case study results of this report highlight the numerical instability of the conventional and stabilized Kalman algorithms. Numerical errors associated with these algorithms can be so large as to obscure important mismodeling effects and thus give misleading estimates of filter accuracy. The positive result of this study is that the Bierman-Thornton U-D covariance factorization algorithm is computationally efficient, with CPU costs that differ negligibly from the conventional Kalman costs. In addition, accuracy of the U-D filter using single-precision arithmetic consistently matches the double-precision reference results. Numerical stability of the U-D filter is further demonstrated by its insensitivity of variations in the a priori statistics.
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.
NASA Technical Reports Server (NTRS)
Luthcke, S. B.; Zelensky, N. P.; Lemoine, Frank G.; Chinn, D. S.; Williams, T. A.
2002-01-01
Jason-1, launched on December 7, 2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. T/P has demonstrated that the time variation of ocean topography can be determined with an accuracy of a few centimeters, thanks to the availability of highly accurate orbits based primarily on SLR+DORIS tracking. The Jason-1 mission is intended to continue measurement of the ocean surface with the same, if not better accuracy. Fortunately, Jason-1 POD can rely on four independent tracking data types available including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. Orbit solutions computed using individual and various combinations of GPS, SLR, DORIS and altimeter crossover data types have been determined from over 100 days of Jason-1 tracking data. The performance of the orbit solutions and tracking data has been evaluated. Orbit solution evaluation and comparison has provided insight into possible areas of refinement. Several aspects of the POD process are examined to obtain orbit improvements including measurement modeling, force modeling and solution strategy. The results of these analyses will be presented.
NASA Technical Reports Server (NTRS)
Luthcke, Scott B.; Zelensky, N. P.; Rowlands, D. D.; Lemoine, F. G.; Chinn, D. S.; Williams, T. A.
2002-01-01
Jason-1, launched on December 7,2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. T P has demonstrated that the time variation of ocean topography can be determined with an accuracy of a few centimeters, thanks to the availability of highly accurate orbits based primarily on SLR+DORIS tracking. The Jason-1 mission is intended to continue measurement of the ocean surface with the same, if not better accuracy. Fortunately, Jason- 1 POD can rely on four independent tracking data types available including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. Orbit solutions computed using individual and various combinations of GPS, SLR, DORIS and altimeter crossover data types have been determined from over 100 days of Jason-1 tracking data, The performance of the orbit solutions and tracking data has been evaluated. Orbit solution evaluation and comparison has provided insight into possible areas of refinement. Several aspects of the POD process are examined to obtain orbit improvements including measurement modeling, force modeling and solution strategy. The results of these analyses will be presented.
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.
20 CFR 405.210 - How to request review of an initial determination.
Code of Federal Regulations, 2010 CFR
2010-04-01
... 20 Employees' Benefits 2 2010-04-01 2010-04-01 false How to request review of an initial determination. 405.210 Section 405.210 Employees' Benefits SOCIAL SECURITY ADMINISTRATION ADMINISTRATIVE REVIEW PROCESS FOR ADJUDICATING INITIAL DISABILITY CLAIMS Review of Initial Determinations by a Federal Reviewing Official § 405.210 How to request...
Code of Federal Regulations, 2010 CFR
2010-10-01
... after the initial one or two months. 233.26 Section 233.26 Public Welfare Regulations Relating to Public... PROGRAMS § 233.26 Retrospective budgeting; determining eligibility after the initial one or two months. (a... specify that eligibility, following the initial one or two months under § 233.24, shall be determined...
40 CFR 179.110 - Determination by Administrator to review initial decision.
Code of Federal Regulations, 2010 CFR
2010-07-01
... review initial decision. 179.110 Section 179.110 Protection of Environment ENVIRONMENTAL PROTECTION... Determination by Administrator to review initial decision. Within 10 days following the expiration of the time..., and serve on the parties, a notice of the Administrator's determination to review the initial...
NASA Astrophysics Data System (ADS)
Zhou, ShanShi; Hu, XiaoGong; Wu, Bin; Liu, Li; Qu, WeiJing; Guo, Rui; He, Feng; Cao, YueLing; Wu, XiaoLi; Zhu, LingFeng; Shi, Xin; Tan, HongLi
2011-06-01
Aiming at regional services, the space segment of COMPASS (Phase I) satellite navigation system is a constellation of Geostationary Earth Orbit (GEO), Inclined Geostationary Earth Orbit (IGSO) and Medium Earth Orbit (MEO) satellites. Precise orbit determination (POD) for the satellites is limited by the geographic distribution of regional tracking stations. Independent time synchronization (TS) system is developed to supplement the regional tracking network, and satellite clock errors and orbit data may be obtained by simultaneously processing both tracking data and TS data. Consequently, inconsistency between tracking system and TS system caused by remaining instrumental errors not calibrated may decrease navigation accuracy. On the other hand, POD for the mixed constellation of GEO/IGSO/MEO with the regional tracking network leads to parameter estimations that are highly correlated. Notorious example of correlation is found between GEO's orbital elements and its clock errors. We estimate orbital elements and clock errors for a 3GEO+2IGSO constellation in this study using a multi-satellite precise orbit determination (MPOD) strategy, with which clock error elimination algorithm is applied to separate orbital and clock estimates to improve numerical efficiency. Satellite Laser Ranging (SLR) data are used to evaluate User Ranging Error (URE), which is the orbital error projected on a receiver's line-of-sight direction. Two-way radio-wave time transfer measurements are used to evaluate clock errors. Experimenting with data from the regional tracking network, we conclude that the fitting of code data is better than 1 m in terms of Root-Mean-Square (RMS), and fitting of carrier phase is better than 1 cm. For orbital evaluation, difference between computed receiver-satellite ranging based on estimated orbits and SLR measurements is better than 1 m (RMS). For clock estimates evaluation, 2-hour linear-fitting shows that the satellite clock rates are about 1.E-10 s
NASA Astrophysics Data System (ADS)
Milley, Ellen Palesa
The physical properties of the meteoroid population were investigated through combining data from a number of fireball camera networks. PE values, as a measure of meteoroid strength, were calculated and linked with other observational criteria (Tisserand parameter, meteor shower identification). The historic divisions for fireball types based on the PE criterion were not observed in the large data set, but a correlation with source region was recognized. Meteor showers demonstrated different amounts of variation in PE values potentially related to the materials found in each parent comet. The trajectory and pre-fall orbit for the Buzzard Coulee meteoroid were determined through the calibration of shadows cast by the fireball. The method of using shadows to triangulate a trajectory was developed and evaluated. The best fit trajectory was coupled with an initial velocity of 18.0 km/s to compute the heliocentric orbit. Buzzard Coulee fell from a modestly inclined near-Earth Apollo orbit. It is the 12th fallen meteorite to be associated with an orbit.
NASA Technical Reports Server (NTRS)
Throckmorton, D. A.
1982-01-01
Temperatures measured at the aerodynamic surface of the Orbiter's thermal protection system (TPS), and calorimeter measurements, are used to determine heating rates to the TPS surface during atmospheric entry. On the Orbiter leeside, where convective heating rates are low, it is possible that a significant portion of the total energy input may result from solar radiation, and for the wing, cross radiation from the hot (relatively) Orbiter fuselage. In order to account for the potential impact of these sources, values of solar- and cross-radiation heat transfer are computed, based upon vehicle trajectory and attitude information and measured surface temperatures. Leeside heat-transfer data from the STS-2 mission are presented, and the significance of solar radiation and fuselage-to-wing cross-radiation contributions to total energy input to Orbiter leeside surfaces is assessed.
Lewis, Karen M.; Fujii, Yuka
2014-08-20
We survey the methods proposed in the literature for detecting moons of extrasolar planets in terms of their ability to distinguish between prograde and retrograde moon orbits, an important tracer of the moon formation channel. We find that most moon detection methods, in particular, sensitive methods for detecting moons of transiting planets, cannot observationally distinguishing prograde and retrograde moon orbits. The prograde and retrograde cases can only be distinguished where the dynamical evolution of the orbit due to, e.g., three body effects is detectable, where one of the two cases is dynamically unstable, or where new observational facilities, which can implement a technique capable of differentiating the two cases, come online. In particular, directly imaged planets are promising targets because repeated spectral and photometric measurements, which are required to determine moon orbit direction, could also be conducted with the primary interest of characterizing the planet itself.
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.
GPS interferometric attitude and heading determination - Initial flight test results
NASA Technical Reports Server (NTRS)
Van Graas, Frank; Braasch, Michael
1992-01-01
Attitude and heading determination using GPS interferometry is a well-understood concept. However, efforts have been concentrated mainly in the development of robust algorithms and applications for low-dynamic, rigid platforms (e.g., shipboard). This paper presents results of what is believed to be the first real-time flight test of a GPS attitude and heading determination system. Signals from four antennas are processed by a 24-channel GPS receiver. Data from the receiver are sent to a microcomputer for storage and further computations. Attitude and heading data are sent to a second computer for display on a software-generated artificial horizon. Demonstration of this technique proves its candidacy for augmentation of aircraft state estimation for flight control and navigation, as well as for numerous other applications.
NASA Astrophysics Data System (ADS)
Peng, D. J.; Wu, B.
2012-01-01
With the availability of precise GPS ephemeris and clock solution, the ionospheric range delay is left as the dominant error sources in the post-processing of space-borne GPS data from single-frequency receivers. Thus, the removal of ionospheric effects is a major prerequisite for an improved orbit reconstruction of LEO satellites equipped with low cost single-frequency GPS receivers. In this paper, the use of Global Ionospheric Maps (GIM) in kinematic and dynamic orbit determination for LEO satellites with single-frequency GPS measurements is discussed first,and then, estimating the scale factor of ionosphere to remove the ionospheric effects in C/A code pseudo-range measurements in both kinematic and adynamia orbit defemination approaches is addressed. As it is known the ionospheric path delay of space-borne GPS signals is strongly dependent on the orbit altitudes of LEO satellites, we selected real space-borne GPS data from CHAMP, GRACE, TerraSAR-X and SAC-C satellites with altitudes between 300 km and 800 km as sample data in this paper. It is demonstrated that the approach of eliminating ionospheric effects in space-borne C/A code pseudo-range by estimating the scale factor of ionosphere is highly effective. Employing this approach, the accuracy of both kinematic and dynamic orbits can be improved notably. Among those five LEO satellites, CHAMP with the lowest orbit altitude has the most remarkable orbit accuracy improvements, which are 55.6% and 47.6% for kinematic and dynamic approaches, respectively. SAC-C with the highest orbit altitude has the least orbit accuracy improvements accordingly, which are 47.8% and 38.2%, respectively.
NASA Astrophysics Data System (ADS)
Hackel, Stefan; Montenbruck, Oliver; Steigenberger, -Peter; Eineder, Michael; Gisinger, Christoph
Remote sensing satellites support a broad range of scientific and commercial applications. The two radar imaging satellites TerraSAR-X and TanDEM-X provide spaceborne Synthetic Aperture Radar (SAR) and interferometric SAR data with a very high accuracy. The increasing demand for precise radar products relies on sophisticated validation methods, which require precise and accurate orbit products. Basically, the precise reconstruction of the satellite’s trajectory is based on the Global Positioning System (GPS) measurements from a geodetic-grade dual-frequency receiver onboard the spacecraft. The Reduced Dynamic Orbit Determination (RDOD) approach utilizes models for the gravitational and non-gravitational forces. Following a proper analysis of the orbit quality, systematics in the orbit products have been identified, which reflect deficits in the non-gravitational force models. A detailed satellite macro model is introduced to describe the geometry and the optical surface properties of the satellite. Two major non-gravitational forces are the direct and the indirect Solar Radiation Pressure (SRP). Due to the dusk-dawn orbit configuration of TerraSAR-X, the satellite is almost constantly illuminated by the Sun. Therefore, the direct SRP has an effect on the lateral stability of the determined orbit. The indirect effect of the solar radiation principally contributes to the Earth Radiation Pressure (ERP). The resulting force depends on the sunlight, which is reflected by the illuminated Earth surface in the visible, and the emission of the Earth body in the infrared spectra. Both components of ERP require Earth models to describe the optical properties of the Earth surface. Therefore, the influence of different Earth models on the orbit quality is assessed within the presentation. The presentation highlights the influence of non-gravitational force and satellite macro models on the orbit quality of TerraSAR-X.
Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods
NASA Technical Reports Server (NTRS)
Khutorovsky, Z.; Kamensky, S.; Sbytov, N.; Alfriend, K. T.
2007-01-01
The problem of determining the orbit of a space object from measurements based on one pass through the field of view of a radar is not a new one. Extensive research in this area has been carried out in the USA and Russia since the late 50s when these countries started the development of ballistic missile defense (BMD) and Early Warning systems. In Russia these investigations got additional stimulation in the early 60s after the decision to create a Space Surveillance System, whose primary task would be the maintenance of the satellite catalog. These problems were the focus of research interest until the middle 70s when the appropriate techniques and software were implemented for all radars. Then for more than 20 years no new research papers appeared on this subject. This produced an impression that all the problems of track determination based on one pass had been solved and there was no need for further research. In the late 90s interest in this problem arose again in relation to the following. It was estimated that there would be greater than 100,000 objects with size greater than 1-2 cm and collision of an operational spacecraft with any of these objects could have catastrophic results. Thus, for prevention of hazardous approaches and collisions with valuable spacecraft the existing satellite catalog should be extended by at least an order of magnitude This is a very difficult scientific and engineering task. One of the issues is the development of data fusion procedures and the software capable of maintaining such a huge catalog in near real time. The number of daily processed measurements (of all types, radar and optical) for such a system may constitute millions, thus increasing the number of measurements by at least an order of magnitude. Since we will have ten times more satellites and measurements the computer effort required for the correlation of measurements will be two orders of magnitude greater. This could create significant problems for processing