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.
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
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 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.
Radio interferometric measurements for accurate planetary orbiter navigation
NASA Technical Reports Server (NTRS)
Poole, S. R.; Ananda, M.; Hildebrand, C. E.
1979-01-01
The use of narrowband delta-VLBI to achieve accurate orbit determination is presented by viewing a spacecraft from widely separated stations followed by viewing a nearby quasar from the same stations. Current analysis is examined that establishes the orbit determination accuracy achieved with data arcs spanning up to 3.5 d. Strategies for improving prediction accuracy are given, and the performance of delta-VLBI is compared with conventional radiometric tracking data. It is found that accuracy 'within the fit' is on the order of 0.5 km for data arcs having delta-VLBI on the ends of the arcs and for arc lengths varying from one baseline to 3.5 d. The technique is discussed with reference to the proposed Venus Orbiting Imaging Radar mission.
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…
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.
Accurate orbit propagation with planetary close encounters
NASA Astrophysics Data System (ADS)
Baù, Giulio; Milani Comparetti, Andrea; Guerra, Francesca
2015-08-01
We tackle the problem of accurately propagating the motion of those small bodies that undergo close approaches with a planet. The literature is lacking on this topic and the reliability of the numerical results is not sufficiently discussed. The high-frequency components of the perturbation generated by a close encounter makes the propagation particularly challenging both from the point of view of the dynamical stability of the formulation and the numerical stability of the integrator. In our approach a fixed step-size and order multistep integrator is combined with a regularized formulation of the perturbed two-body problem. When the propagated object enters the region of influence of a celestial body, the latter becomes the new primary body of attraction. Moreover, the formulation and the step-size will also be changed if necessary. We present: 1) the restarter procedure applied to the multistep integrator whenever the primary body is changed; 2) new analytical formulae for setting the step-size (given the order of the multistep, formulation and initial osculating orbit) in order to control the accumulation of the local truncation error and guarantee the numerical stability during the propagation; 3) a new definition of the region of influence in the phase space. We test the propagator with some real asteroids subject to the gravitational attraction of the planets, the Yarkovsky and relativistic perturbations. Our goal is to show that the proposed approach improves the performance of both the propagator implemented in the OrbFit software package (which is currently used by the NEODyS service) and of the propagator represented by a variable step-size and order multistep method combined with Cowell's formulation (i.e. direct integration of position and velocity in either the physical or a fictitious time).
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.
Extremely Accurate On-Orbit Position Accuracy using TDRSS
NASA Technical Reports Server (NTRS)
Stocklin, Frank; Toral, Marco; Bar-Sever, Yoaz; Rush, John
2006-01-01
NASA is planning to launch a new service for Earth satellites providing them with precise GPS differential corrections and other ancillary information enabling decimeter level orbit determination accuracy and nanosecond time-transfer accuracy, onboard, in real-time. The TDRSS Augmentation Service for Satellites (TASS) will broadcast its message on the S-band multiple access forward channel of NASA s Tracking and Data Relay Satellite System (TDRSS). The satellite's phase array antenna has been configured to provide a wide beam, extending coverage up to 1000 km altitude over the poles. Global coverage will be ensured with broadcast from three or more TDRSS satellites. The GPS differential corrections are provided by the NASA Global Differential GPS (GDGPS) System, developed and operated by JPL. The GDGPS System employs global ground network of more than 70 GPS receivers to monitor the GPS constellation in real time. The system provides real-time estimates of the GPS satellite states, as well as many other real-time products such as differential corrections, global ionospheric maps, and integrity monitoring. The unique multiply redundant architecture of the GDGPS System ensures very high reliability, with 99.999% demonstrated since the inception of the system in early 2000. The estimated real time GPS orbit and clock states provided by the GDGPS system are accurate to better than 20 cm 3D RMS, and have been demonstrated to support sub-decimeter real time positioning and orbit determination for a variety of terrestrial, airborne, and spaceborne applications. In addition to the GPS differential corrections, TASS will provide real-time Earth orientation and solar flux information that enable precise onboard knowledge of the Earth-fixed position of the spacecraft, and precise orbit prediction and planning capabilities. TASS will also provide 5 seconds alarms for GPS integrity failures based on the unique GPS integrity monitoring service of the GDGPS System.
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.
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.
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 Technical Reports Server (NTRS)
Tang, Charles C. H.
1988-01-01
By using Von Zeipel's generating function procedure the perturbing earth gravitational potential is averaged with respect to the fast variable (mean anomaly) and a set of 'fictitous' mean orbital elements which can be used as a long-term satellite orbit predictor is obtained. The set of elements is shown to be a function of the nonlinear square of the second zonal harmonic coefficient. It is found that the long-term orbit prediction using the 'fictitous' mean elements is as accurate as that using the osculating elements, but has a computing speed about two orders of magnitude faster. For short-term orbit predictions, the osculating elements approach must be used.
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.
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.
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.
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.
NASA Astrophysics Data System (ADS)
Iorio, L.
2016-01-01
By using the most recently published Doppler tomography measurements and accurate theoretical modelling of the oblateness-driven orbital precessions, we tightly constrain some of the physical and orbital parameters of the planetary system hosted by the fast rotating star WASP-33. In particular, the measurements of the orbital inclination ip to the plane of the sky and of the sky-projected spin-orbit misalignment λ at two epochs about six years apart allowed for the determination of the longitude of the ascending node Ω and of the orbital inclination I to the apparent equatorial plane at the same epochs. As a consequence, average rates of change dot{Ω }_exp, dot{I}_exp of this two orbital elements, accurate to a ≈10-2 deg yr-1 level, were calculated as well. By comparing them to general theoretical expressions dot{Ω }_{J_2}, dot{I}_{J_2} for their precessions induced by an oblate star whose symmetry axis is arbitrarily oriented, we were able to determine the angle i⋆ between the line of sight the star's spin {S}^{star } and its first even zonal harmonic J_2^{star } obtaining i^{star } = {142}^{+10}_{-11} deg, J_2^{star } = 2.1^{+0.8}_{-0.5}times; 10^{-4}. As a by-product, the angle between {S}^{star } and the orbital angular momentum L is as large as about ψ ≈ 100 ° psi; ^{2008} = 99^{+5}_{-4} deg, ψ ^{{2014}} = 103^{+5}_{-4} deg and changes at a rate dot{ψ }= 0.{7}^{+1.5}_{-1.6} deg {yr}^{-1}. The predicted general relativistic Lense-Thirring precessions, of the order of ≈10-3deg yr-1, are, at present, about one order of magnitude below the measurability threshold.
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.
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.
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.
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).
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.
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.
Accurate determination of characteristic relative permeability curves
NASA Astrophysics Data System (ADS)
Krause, Michael H.; Benson, Sally M.
2015-09-01
A recently developed technique to accurately characterize sub-core scale heterogeneity is applied to investigate the factors responsible for flowrate-dependent effective relative permeability curves measured on core samples in the laboratory. The dependency of laboratory measured relative permeability on flowrate has long been both supported and challenged by a number of investigators. Studies have shown that this apparent flowrate dependency is a result of both sub-core scale heterogeneity and outlet boundary effects. However this has only been demonstrated numerically for highly simplified models of porous media. In this paper, flowrate dependency of effective relative permeability is demonstrated using two rock cores, a Berea Sandstone and a heterogeneous sandstone from the Otway Basin Pilot Project in Australia. Numerical simulations of steady-state coreflooding experiments are conducted at a number of injection rates using a single set of input characteristic relative permeability curves. Effective relative permeability is then calculated from the simulation data using standard interpretation methods for calculating relative permeability from steady-state tests. Results show that simplified approaches may be used to determine flowrate-independent characteristic relative permeability provided flow rate is sufficiently high, and the core heterogeneity is relatively low. It is also shown that characteristic relative permeability can be determined at any typical flowrate, and even for geologically complex models, when using accurate three-dimensional models.
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.
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.
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.
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.
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.
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.
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 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
measurements that would be needed to meet the required orbit determination accuracies. Analysts used the Orbit Determination Error Analysis System (ODEAS) to perform covariance analyses using various tracking data schedules. From this analysis, it was determined that 3.5 hours of DSN TRK-2-34 range and Doppler tracking data every other day would suffice to meet the predictive orbit knowledge accuracies in the Lissajous region. The results of this analysis are presented. Both GTDS and ODTK have high-fidelity environmental orbit force models that allow for very accurate orbit estimation in the lunar Lissajous regime. These models include solar radiation pressure, Earth and Moon gravity models, third body gravitational effects from the Sun, and to a lesser extent third body gravitational effects from Jupiter, Venus, Saturn, and Mars. Increased position and velocity uncertainties following each maneuver, due to small execution performance errors, requires that several days of post-maneuver tracking data be processed to converge on an accurate post-maneuver orbit solution. The effects of maneuvers on orbit determination accuracy will be presented, including a comparison of the batch least squares technique to the extended Kalman filter/smoother technique. We will present the maneuver calibration results derived from processing post-maneuver tracking data. A dominant error in the orbit estimation process is the uncertainty in solar radiation pressure and the resultant force on the spacecraft. An estimation of this value can include many related factors, such as the uncertainty in spacecraft reflectivity and surface area which is a function of spacecraft orientation (spin-axis attitude), uncertainty in spacecraft wet mass, and potential seasonal variability due to the changing direction of the Sun line relative to the Earth-Moon Lissajous reference frame. In addition, each spacecraft occasionally enters into Earth or Moon penumbra or umbra and these shadow crossings reduche solar
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.
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.
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.
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.
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.
Accurate abundance determinations in S stars
NASA Astrophysics Data System (ADS)
Neyskens, P.; Van Eck, S.; Plez, B.; Goriely, S.; Siess, L.; Jorissen, A.
2011-12-01
S-type stars are thought to be the first objects, during their evolution on the asymptotic giant branch (AGB), to experience s-process nucleosynthesis and third dredge-ups, and therefore to exhibit s-process signatures in their atmospheres. Until present, the modeling of these processes is subject to large uncertainties. Precise abundance determinations in S stars are of extreme importance for constraining e.g., the depth and the formation of the 13C pocket. In this paper a large grid of MARCS model atmospheres for S stars is used to derive precise abundances of key s-process elements and iron. A first estimation of the atmospheric parameters is obtained using a set of well-chosen photometric and spectroscopic indices for selecting the best model atmosphere of each S star. Abundances are derived from spectral line synthesis, using the selected model atmosphere. Special interest is paid to technetium, an element without stable isotopes. Its detection in stars is considered as the best possible signature that the star effectively populates the thermally-pulsing AGB (TP-AGB) phase of evolution. The derived Tc/Zr abundances are compared, as a function of the derived [Zr/Fe] overabundances, with AGB stellar model predictions. The computed [Zr/Fe] overabundances are in good agreement with the AGB stellar evolution model predictions, while the Tc/Zr abundances are slightly over-predicted. This discrepancy can help to set stronger constraints on nucleosynthesis and mixing mechanisms in AGB stars.
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.
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.
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.
Effective Echo Detection and Accurate Orbit Estimation Algorithms for Space Debris Radar
NASA Astrophysics Data System (ADS)
Isoda, Kentaro; Sakamoto, Takuya; Sato, Toru
Orbit estimation of space debris, objects of no inherent value orbiting the earth, is a task that is important for avoiding collisions with spacecraft. The Kamisaibara Spaceguard Center radar system was built in 2004 as the first radar facility in Japan devoted to the observation of space debris. In order to detect the smaller debris, coherent integration is effective in improving SNR (Signal-to-Noise Ratio). However, it is difficult to apply coherent integration to real data because the motions of the targets are unknown. An effective algorithm is proposed for echo detection and orbit estimation of the faint echoes from space debris. The characteristics of the evaluation function are utilized by the algorithm. Experiments show the proposed algorithm improves SNR by 8.32dB and enables estimation of orbital parameters accurately to allow for re-tracking with a single radar.
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.
Orbit determination methods in view of the PODET project
NASA Astrophysics Data System (ADS)
Deleflie, F.; Coulot, D.; Decosta, R.; Richard, P.
2013-11-01
We present an orbit determination method based on genetic algorithms. Contrary to usual estimation methods mainly based on least-squares methods, these algorithms do not require any a priori knowledge of the initial state vector to be estimated. These algorithms can be applied when a new satellite is launched or for uncatalogued objects We show in this paper preliminary results obtained from an SLR satellite, for which tracking data acquired by the ILRS network enable to build accurate orbital arcs at a few centimeter level, which can be used as a reference orbit. The method is carried out in several steps: (i) an analytical propagation of the equations of motion, (ii) an estimation kernel based on genetic algorithms, which follows the usual steps of such approaches: initialization and evolution of a selected population, so as to determine the best parameters. Each parameter to be estimated, namely each initial keplerian element, has to be searched among an interval that is preliminary chosen.
Meteor orbit determination with improved accuracy
NASA Astrophysics Data System (ADS)
Dmitriev, Vasily; Lupovla, Valery; Gritsevich, Maria
2015-08-01
Modern observational techniques make it possible to retrive meteor trajectory and its velocity with high accuracy. There has been a rapid rise in high quality observational data accumulating yearly. This fact creates new challenges for solving the problem of meteor orbit determination. Currently, traditional technique based on including corrections to zenith distance and apparent velocity using well-known Schiaparelli formula is widely used. Alternative approach relies on meteoroid trajectory correction using numerical integration of equation of motion (Clark & Wiegert, 2011; Zuluaga et al., 2013). In our work we suggest technique of meteor orbit determination based on strict coordinate transformation and integration of differential equation of motion. We demonstrate advantage of this method in comparison with traditional technique. We provide results of calculations by different methods for real, recently occurred fireballs, as well as for simulated cases with a priori known retrieval parameters. Simulated data were used to demonstrate the condition, when application of more complex technique is necessary. It was found, that for several low velocity meteoroids application of traditional technique may lead to dramatically delusion of orbit precision (first of all, due to errors in 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.
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].
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.
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.
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.
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.
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.
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.
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.
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
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.
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).
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.
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 and Accurate Density Determination of Explosives Using Hydrostatic Weighing
B. Olinger
2005-07-01
Precise and accurate density determination requires weight measurements in air and water using sufficiently precise analytical balances, knowledge of the densities of air and water, knowledge of thermal expansions, availability of a density standard, and a method to estimate the time to achieve thermal equilibrium with water. Density distributions in pressed explosives are inferred from the densities of elements from a central slice.
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.
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.
Accurate Energies and Orbital Description in Semi-Local Kohn-Sham DFT
NASA Astrophysics Data System (ADS)
Lindmaa, Alexander; Kuemmel, Stephan; Armiento, Rickard
2015-03-01
We present our progress on a scheme in semi-local Kohn-Sham density-functional theory (KS-DFT) for improving the orbital description while still retaining the level of accuracy of the usual semi-local exchange-correlation (xc) functionals. DFT is a widely used tool for first-principles calculations of properties of materials. A given task normally requires a balance of accuracy and computational cost, which is well achieved with semi-local DFT. However, commonly used semi-local xc functionals have important shortcomings which often can be attributed to features of the corresponding xc potential. One shortcoming is an overly delocalized representation of localized orbitals. Recently a semi-local GGA-type xc functional was constructed to address these issues, however, it has the trade-off of lower accuracy of the total energy. We discuss the source of this error in terms of a surplus energy contribution in the functional that needs to be accounted for, and offer a remedy for this issue which formally stays within KS-DFT, and, which does not harshly increase the computational effort. The end result is a scheme that combines accurate total energies (e.g., relaxed geometries) with an improved orbital description (e.g., improved band structure).
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.
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.
Accurate Method for Determining Adhesion of Cantilever Beams
Michalske, T.A.; de Boer, M.P.
1999-01-08
Using surface micromachined samples, we demonstrate the accurate measurement of cantilever beam adhesion by using test structures which are adhered over long attachment lengths. We show that this configuration has a deep energy well, such that a fracture equilibrium is easily reached. When compared to the commonly used method of determining the shortest attached beam, the present method is much less sensitive to variations in surface topography or to details of capillary drying.
Accurate method for determining adhesion of cantilever beams
de Boer, M.P.; Michalske, T.A.
1999-07-01
Using surface micromachined samples, we demonstrate the accurate measurement of cantilever beam adhesion by using test structures which are adhered over long attachment lengths. We show that this configuration has a deep energy well, such that a fracture equilibrium is easily reached. When compared to the commonly used method of determining the shortest attached beam, the present method is much less sensitive to variations in surface topography or to details of capillary drying. {copyright} {ital 1999 American Institute of Physics.}
2012-01-01
A natural bond orbital (NBO) analysis of unpaired electron spin density in metalloproteins is presented, which allows a fast and robust calculation of paramagnetic NMR parameters. Approximately 90% of the unpaired electron spin density occupies metal–ligand NBOs, allowing the majority of the density to be modeled by only a few NBOs that reflect the chemical bonding environment. We show that the paramagnetic relaxation rate of protons can be calculated accurately using only the metal–ligand NBOs and that these rates are in good agreement with corresponding rates measured experimentally. This holds, in particular, for protons of ligand residues where the point-dipole approximation breaks down. To describe the paramagnetic relaxation of heavy nuclei, also the electron spin density in the local orbitals must be taken into account. Geometric distance restraints for 15N can be derived from the paramagnetic relaxation enhancement and the Fermi contact shift when local NBOs are included in the analysis. Thus, the NBO approach allows us to include experimental paramagnetic NMR parameters of 15N nuclei as restraints in a structure optimization protocol. We performed a molecular dynamics simulation and structure determination of oxidized rubredoxin using the experimentally obtained paramagnetic NMR parameters of 15N. The corresponding structures obtained are in good agreement with the crystal structure of rubredoxin. Thus, the NBO approach allows an accurate description of the geometric structure and the dynamics of metalloproteins, when NMR parameters are available of nuclei in the immediate vicinity of the metal-site. PMID:22329704
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].
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.
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.
Orbital Advection by Interpolation: A Fast and Accurate Numerical Scheme for Super-Fast MHD Flows
Johnson, B M; Guan, X; Gammie, F
2008-04-11
In numerical models of thin astrophysical disks that use an Eulerian scheme, gas orbits supersonically through a fixed grid. As a result the timestep is sharply limited by the Courant condition. Also, because the mean flow speed with respect to the grid varies with position, the truncation error varies systematically with position. For hydrodynamic (unmagnetized) disks an algorithm called FARGO has been developed that advects the gas along its mean orbit using a separate interpolation substep. This relaxes the constraint imposed by the Courant condition, which now depends only on the peculiar velocity of the gas, and results in a truncation error that is more nearly independent of position. This paper describes a FARGO-like algorithm suitable for evolving magnetized disks. Our method is second order accurate on a smooth flow and preserves {del} {center_dot} B = 0 to machine precision. The main restriction is that B must be discretized on a staggered mesh. We give a detailed description of an implementation of the code and demonstrate that it produces the expected results on linear and nonlinear problems. We also point out how the scheme might be generalized to make the integration of other supersonic/super-fast flows more efficient. Although our scheme reduces the variation of truncation error with position, it does not eliminate it. We show that the residual position dependence leads to characteristic radial variations in the density over long integrations.
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.
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.
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 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
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.
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
Laser Ranging for Effective and Accurate Tracking of Space Debris in Low Earth Orbits
NASA Astrophysics Data System (ADS)
Blanchet, Guillaume; Haag, Herve; Hennegrave, Laurent; Assemat, Francois; Vial, Sophie; Samain, Etienne
2013-08-01
The paper presents the results of preliminary design options for an operational laser ranging system adapted to the measurement of the distance of space debris. Thorough analysis of the operational parameters is provided with identification of performance drivers and assessment of enabling design options. Results from performance simulation demonstrate how the range measurement enables improvement of the orbit determination when combined with astrometry. Besides, experimental results on rocket-stage class debris in LEO were obtained by Astrium beginning of 2012, in collaboration with the Observatoire de la Côte d'Azur (OCA), by operating an experimental laser ranging system supported by the MéO (Métrologie Optique) telescope.
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.
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.
Reverse radiance: a fast accurate method for determining luminance
NASA Astrophysics Data System (ADS)
Moore, Kenneth E.; Rykowski, Ronald F.; Gangadhara, Sanjay
2012-10-01
Reverse ray tracing from a region of interest backward to the source has long been proposed as an efficient method of determining luminous flux. The idea is to trace rays only from where the final flux needs to be known back to the source, rather than tracing in the forward direction from the source outward to see where the light goes. Once the reverse ray reaches the source, the radiance the equivalent forward ray would have represented is determined and the resulting flux computed. Although reverse ray tracing is conceptually simple, the method critically depends upon an accurate source model in both the near and far field. An overly simplified source model, such as an ideal Lambertian surface substantially detracts from the accuracy and thus benefit of the method. This paper will introduce an improved method of reverse ray tracing that we call Reverse Radiance that avoids assumptions about the source properties. The new method uses measured data from a Source Imaging Goniometer (SIG) that simultaneously measures near and far field luminous data. Incorporating this data into a fast reverse ray tracing integration method yields fast, accurate data for a wide variety of illumination problems.
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%.
Accurate determination of cobalt traces in several biological reference materials.
Dybczyński, R; Danko, B
1994-01-01
A newly devised, very accurate ("definitive") method for the determination of trace amounts of cobalt in biological materials was validated by the analysis of several certified reference materials. The method is based on a combination of neutron activation and selective and quantitative postirradiation isolation of radiocobalt from practically all other radionuclides by ion-exchange and extraction chromatography followed by gamma-ray spectrometric measurement. The significance of criteria that should be fulfilled in order to accept a given result as obtained by the "definitive method" is emphasized. In view of the demonstrated very good accuracy of the method, it is suggested that our values for cobalt content in those reference materials in which it was originally not certified (SRM 1570 spinach, SRM 1571 orchard leaves, SRM 1577 bovine liver, and Czechoslovak bovine liver 12-02-01) might be used as provisional certified values. PMID:7710879
Fast and accurate automated cell boundary determination for fluorescence microscopy
NASA Astrophysics Data System (ADS)
Arce, Stephen Hugo; Wu, Pei-Hsun; Tseng, Yiider
2013-07-01
Detailed measurement of cell phenotype information from digital fluorescence images has the potential to greatly advance biomedicine in various disciplines such as patient diagnostics or drug screening. Yet, the complexity of cell conformations presents a major barrier preventing effective determination of cell boundaries, and introduces measurement error that propagates throughout subsequent assessment of cellular parameters and statistical analysis. State-of-the-art image segmentation techniques that require user-interaction, prolonged computation time and specialized training cannot adequately provide the support for high content platforms, which often sacrifice resolution to foster the speedy collection of massive amounts of cellular data. This work introduces a strategy that allows us to rapidly obtain accurate cell boundaries from digital fluorescent images in an automated format. Hence, this new method has broad applicability to promote biotechnology.
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.
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.
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.
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.
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.
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.
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.
Nanoparticle Counting: Towards Accurate Determination of the Molar Concentration
Shang, Jing; Gao, Xiaohu
2014-01-01
Summary Innovations in nanotechnology have brought tremendous opportunities for the advancement of many research frontiers, ranging from electronics, photonics, energy, to medicine. To maximize the benefits of nano-scaled materials in different devices and systems, precise control of their concentration is a prerequisite. While concentrations of nanoparticles have been provided in other forms (e.g., mass), accurate determination of molar concentration, arguably the most useful one for chemical reactions and applications, has been a major challenge (especially for nanoparticles smaller than 30 nm). Towards this significant yet chronic problem, a variety of strategies are currently under development. Most of these strategies are applicable to a specialized group of nanoparticles due to their restrictions on the composition and size ranges of nanoparticles. As research and uses of nanomaterials being explored in an unprecedented speed, it is necessary to develop universal strategies that are easy to use, and compatible with nanoparticles of different sizes, compositions, and shapes. This review outlines the theories and applications of current strategies to measure nanoparticle molar concentration, discusses the advantages and limitations of these methods, and provides insights into future directions. PMID:25099190
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.
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)
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.
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
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.
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.
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.
Fast Geometric Method for Calculating Accurate Minimum Orbit Intersection Distances (MOIDs)
NASA Astrophysics Data System (ADS)
Wiźniowski, T.; Rickman, H.
2013-06-01
We present a new method to compute Minimum Orbit Intersection Distances (MOIDs) for arbitrary pairs of heliocentric orbits and compare it with Giovanni Gronchi's algebraic method. Our procedure is numerical and iterative, and the MOID configuration is found by geometric scanning and tuning. A basic element is the meridional plane, used for initial scanning, which contains one of the objects and is perpendicular to the orbital plane of the other. Our method also relies on an efficient tuning technique in order to zoom in on the MOID configuration, starting from the first approximation found by scanning. We work with high accuracy and take special care to avoid the risk of missing the MOID, which is inherent to our type of approach. We demonstrate that our method is both fast, reliable and flexible. It is freely available and its source Fortran code downloadable via our web page.
Optimum satellite orbits for accurate measurement of the earth's radiation budget, summary
NASA Technical Reports Server (NTRS)
Campbell, G. G.; Vonderhaar, T. H.
1978-01-01
The optimum set of orbit inclinations for the measurement of the earth radiation budget from spacially integrating sensor systems was estimated for two and three satellite systems. The best set of the two were satellites at orbit inclinations of 80 deg and 50 deg; of three the inclinations were 80 deg, 60 deg and 50 deg. These were chosen on the basis of a simulation of flat plate and spherical detectors flying over a daily varying earth radiation field as measured by the Nimbus 3 medium resolution scanners. A diurnal oscillation was also included in the emitted flux and albedo to give a source field as realistic as possible. Twenty three satellites with different inclinations and equator crossings were simulated, allowing the results of thousand of multisatellite sets to be intercompared. All were circular orbits of radius 7178 kilometers.
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.
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.
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.
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.
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 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)
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.
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.
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
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.
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).
Accurately Determining the Risks of Rising Sea Level
NASA Astrophysics Data System (ADS)
Marbaix, Philippe; Nicholls, Robert J.
2007-10-01
With the highest density of people and the greatest concentration of economic activity located in the coastal regions, sea level rise is an important concern as the climate continues to warm. Subsequent flooding may potentially disrupt industries, populations, and livelihoods, particularly in the long term if the climate is not quickly stabilized [McGranahan et al., 2007; Tol et al., 2006]. To help policy makers understand these risks, a more accurate description of hazards posed by rising sea levels is needed at the global scale, even though the impacts in specific regions are better known.
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.
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.
Accurate Determination of Membrane Dynamics with Line-Scan FCS
Ries, Jonas; Chiantia, Salvatore; Schwille, Petra
2009-01-01
Here we present an efficient implementation of line-scan fluorescence correlation spectroscopy (i.e., one-dimensional spatio-temporal image correlation spectroscopy) using a commercial laser scanning microscope, which allows the accurate measurement of diffusion coefficients and concentrations in biological lipid membranes within seconds. Line-scan fluorescence correlation spectroscopy is a calibration-free technique. Therefore, it is insensitive to optical artifacts, saturation, or incorrect positioning of the laser focus. In addition, it is virtually unaffected by photobleaching. Correction schemes for residual inhomogeneities and depletion of fluorophores due to photobleaching extend the applicability of line-scan fluorescence correlation spectroscopy to more demanding systems. This technique enabled us to measure accurate diffusion coefficients and partition coefficients of fluorescent lipids in phase-separating supported bilayers of three commonly used raft-mimicking compositions. Furthermore, we probed the temperature dependence of the diffusion coefficient in several model membranes, and in human embryonic kidney cell membranes not affected by temperature-induced optical aberrations. PMID:19254560
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
Accurate Mass Determinations in Decay Chains with Missing Energy
Cheng, H.-C; Gunion, John F.; Han Zhenyu; Engelhardt, Dalit; McElrath, Bob
2008-06-27
Many beyond the standard model theories include a stable dark matter candidate that yields missing or invisible energy in collider detectors. If observed at the CERN Large Hadron Collider, we must determine if its mass and other properties (and those of its partners) predict the correct dark matter relic density. We give a new procedure for determining its mass with small error.
Accurate mass determinations in decay chains with missing energy.
Cheng, Hsin-Chia; Engelhardt, Dalit; Gunion, John F; Han, Zhenyu; McElrath, Bob
2008-06-27
Many beyond the standard model theories include a stable dark matter candidate that yields missing or invisible energy in collider detectors. If observed at the CERN Large Hadron Collider, we must determine if its mass and other properties (and those of its partners) predict the correct dark matter relic density. We give a new procedure for determining its mass with small error. PMID:18643654
Determination of sockage for accurate rough rice quality assessment
Technology Transfer Automated Retrieval System (TEKTRAN)
Determination of dockage of freshly harvested rice is crucial for precise development of a universal rice shrinking chart. The objectives of this research were to determine the effect of different factors, including rice variety, farm location, harvest moisture and time, drying, dropping, weather ev...
Accurate Determination of Conformational Transitions in Oligomeric Membrane Proteins
Sanz-Hernández, Máximo; Vostrikov, Vitaly V.; Veglia, Gianluigi; De Simone, Alfonso
2016-01-01
The structural dynamics governing collective motions in oligomeric membrane proteins play key roles in vital biomolecular processes at cellular membranes. In this study, we present a structural refinement approach that combines solid-state NMR experiments and molecular simulations to accurately describe concerted conformational transitions identifying the overall structural, dynamical, and topological states of oligomeric membrane proteins. The accuracy of the structural ensembles generated with this method is shown to reach the statistical error limit, and is further demonstrated by correctly reproducing orthogonal NMR data. We demonstrate the accuracy of this approach by characterising the pentameric state of phospholamban, a key player in the regulation of calcium uptake in the sarcoplasmic reticulum, and by probing its dynamical activation upon phosphorylation. Our results underline the importance of using an ensemble approach to characterise the conformational transitions that are often responsible for the biological function of oligomeric membrane protein states. PMID:26975211
Efficient determination of accurate atomic polarizabilities for polarizeable embedding calculations.
Schröder, Heiner; Schwabe, Tobias
2016-08-15
We evaluate embedding potentials, obtained via various methods, used for polarizable embedding computations of excitation energies of para-nitroaniline in water and organic solvents as well as of the green fluorescent protein. We found that isotropic polarizabilities derived from DFTD3 dispersion coefficients correlate well with those obtained via the LoProp method. We show that these polarizabilities in conjunction with appropriately derived point charges are in good agreement with calculations employing static multipole moments up to quadrupoles and anisotropic polarizabilities for both computed systems. The (partial) use of these easily-accessible parameters drastically reduces the computational effort to obtain accurate embedding potentials especially for proteins. © 2016 The Authors. Journal of Computational Chemistry Published by Wiley Periodicals, Inc. PMID:27317509
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
Apparatus enables accurate determination of alkali oxides in alkali metals
NASA Technical Reports Server (NTRS)
Dupraw, W. A.; Gahn, R. F.; Graab, J. W.; Maple, W. E.; Rosenblum, L.
1966-01-01
Evacuated apparatus determines the alkali oxide content of an alkali metal by separating the metal from the oxide by amalgamation with mercury. The apparatus prevents oxygen and moisture from inadvertently entering the system during the sampling and analytical procedure.
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
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.
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.
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.
Applications of accurate isentropic exponent determination for fuel gas measurement
Pack, D.J.; Edwards, T.J.; Fawcett, D.
1996-07-01
This paper discusses the determination and application of the isentropic exponent to the various thermodynamic processes found in a high-pressure natural gas transmission system. Increasing demands for more precise measurement of natural gas, coupled with the need for greater efficiency and accountability of transportation and processing operations, had led to the research and development of gas thermodynamic properties including isentropic exponent. The isentropic exponent has many applications, some of which include: the determination of the expansion factor {epsilon}, for calculation of flow using an orifice or venturi-type meter; the volumetric efficiency in a reciprocating compressor; the determination of the compression head for a centrifugal compressor; the engine power required for the given conditions for a gas compressor; the calculation of discharge temperatures for compressors; and the direct measurement of gas density. As can be appreciated, the application of an incorrect value for the isentropic exponent represents an error in the parameter determined. For large volume gas flows, this can translate into a significant cost penalty.
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
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.
The Greenhouse Effect - Determination From Accurate Surface Longwave Radiation Measurements
NASA Astrophysics Data System (ADS)
Philipona, R.
Longwave radiation measurements have been drastically improved in recent years. Uncertainty levels down to s2 Wm-2 are realistic and achieved during long-term ´ longwave irradiance measurements. Longwave downward irradiance measurements together with temperature and humidity measurements at the station are used to sepa- rate clear-sky from cloudy-sky situations. Longwave net radiation separated between clear-sky and all-sky situations allows to determine the longwave cloud radiative forc- ing at the station. For clear-sky situations radiative transfer models demonstrate a lin- ear relation between longwave downward radiation and the greenhouse radiative flux. Clear-sky longwave radiation, temperature and humidity for different atmospheres and different altitudes were modeled with the MODTRAN radiative transfer code and compared to longwave radiation, temperature and humidity measured at 4 radiation stations of the Alpine Surface Radiation Budget (ASRB) network at similar altitude and with corresponding atmospheres. At the 11 ASRB stations the clear-sky green- house effect was determined by using clear-sky longwave downward measurements and MODTRAN model calculations. The all-sky greenhouse effect was determined by adding the longwave cloud radiative forcing to the clear-sky greenhouse radiative flux. The altitude dependence of annual and seasonal mean values of the greenhouse effect will be shown for the altitude range of 400 to 3600 meter a.s.l. in the Alps.
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.
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.
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.
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.
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
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 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.
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.
Accurate Optical Target Pose Determination for Applications in Aerial Photogrammetry
NASA Astrophysics Data System (ADS)
Cucci, D. A.
2016-06-01
We propose a new design for an optical coded target based on concentric circles and a position and orientation determination algorithm optimized for high distances compared to the target size. If two ellipses are fitted on the edge pixels corresponding to the outer and inner circles, quasi-analytical methods are known to obtain the coordinates of the projection of the circles center. We show the limits of these methods for quasi-frontal target orientations and in presence of noise and we propose an iterative refinement algorithm based on a geometric invariant. Next, we introduce a closed form, computationally inexpensive, solution to obtain the target position and orientation given the projected circle center and the parameters of the outer circle projection. The viability of the approach is demonstrated based on aerial pictures taken by an UAV from elevations between 10 to 100 m. We obtain a distance RMS below 0.25 % under 50 m and below 1 % under 100 m with a target size of 90 cm, part of which is a deterministic bias introduced by image exposure.
An autonomous orbit determination method for MEO and LEO satellite
NASA Astrophysics Data System (ADS)
Zhang, Hui; Wang, Jin; Yu, Guobin; Zhong, Jie; Lin, Ling
2014-09-01
A reliable and secure navigation system and assured autonomous capability of satellite are in high demand in case of emergencies in space. This paper introduces a novel autonomous orbit determination method for Middle-Earth-Orbit and Low-Earth-Orbit (MEO and LEO) satellite by observing space objects whose orbits are known. Generally, the geodetic satellites, such as LAGEOS and ETALONS, can be selected as the space objects here. The precision CCD camera on tracking gimbal can make a series of photos of the objects and surrounding stars when MEO and LEO satellite encounters the space objects. Then the information processor processes images and attains sightings and angular observations of space objects. Several clusters of such angular observations are incorporated into a batch least squares filter to obtain an orbit determination solution. This paper describes basic principle and builds integrated mathematical model. The accuracy of this method is analyzed by means of computer simulation. Then a simulant experiment system is built, and the experimental results demonstrate the feasibility and effectiveness of this method. The experimental results show that this method can attain the accuracy of 150 meters with angular observations of 1 arcsecond system error.
Mars Science Laboratory Orbit Determination Data Pre-Processing
NASA Technical Reports Server (NTRS)
Gustafson, Eric D.; Kruizinga, Gerhard L.; Martin-Mur, Tomas J.
2013-01-01
The Mars Science Laboratory (MSL) was spin-stabilized during its cruise to Mars. We discuss the effects of spin on the radiometric data and how the orbit determination team dealt with them. Additionally, we will discuss the unplanned benefits of detailed spin modeling including attitude estimation and spacecraft clock correlation.
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.
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.
Minesaki, Yukitaka
2013-08-01
For the restricted three-body problem, we propose an accurate orbital integration scheme that retains all conserved quantities of the two-body problem with two primaries and approximately preserves the Jacobi integral. The scheme is obtained by taking the limit as mass approaches zero in the discrete-time general three-body problem. For a long time interval, the proposed scheme precisely reproduces various periodic orbits that cannot be accurately computed by other generic integrators.
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.
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
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.
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.
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.
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.
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.
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
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.
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.
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.
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.
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.
Accurate determination of the vapor pressure of potassium using optical absorption
NASA Technical Reports Server (NTRS)
Shirinzadeh, B.; Wang, C. C.
1983-01-01
The vapor pressure of potassium has been measured in absorption using a CW tunable laser and calibrated against the accurate radiative lifetime of the 4s-4p doublet of potassium. An accurate value of 20,850 + or - 30 cal/mol for the heat of vaporization (from the liquid phase) at the melting point was determined.
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.
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.
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.
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.
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.
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.
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
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
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.
Fast and accurate determination of the Wigner rotation matrices in the fast multipole method.
Dachsel, Holger
2006-04-14
In the rotation based fast multipole method the accurate determination of the Wigner rotation matrices is essential. The combination of two recurrence relations and the control of the error accumulations allow a very precise determination of the Wigner rotation matrices. The recurrence formulas are simple, efficient, and numerically stable. The advantages over other recursions are documented. PMID:16626188
Technology Transfer Automated Retrieval System (TEKTRAN)
Analytical methods for the determination of mycotoxins in foods are commonly based on chromatographic techniques (GC, HPLC or LC-MS). Although these methods permit a sensitive and accurate determination of the analyte, they require skilled personnel and are time-consuming, expensive, and unsuitable ...
Accurate potential energy functions, non-adiabatic and spin-orbit couplings in the ZnH+ system
NASA Astrophysics Data System (ADS)
Liang, Guiying; Liu, Xiaoting; Zhang, Xiaomei; Xu, Haifeng; Yan, Bing
2016-03-01
A high-level ab initio calculation on the ZnH+ cation has been carried out with the multi-reference configuration interaction method plus Davison correction (MRCI + Q). The scalar relativistic effect is included by using the Douglas-Kroll-Hess (DKH) method. The calculated potential energy curves (PECs) of the 7 Λ-S states are associated with the dissociation limits of Zn+(2Sg) + H(2Sg), Zn(1Sg) + H+(1Sg), and Zn+(2Pu) + H(2Sg), respectively (The Λ-S state is labeled as 2S + 1Λ, in which Λ is the quantum number for the projection along the internuclear axis of the total electronic orbital angular momentum and S is the total electron spin). The spectroscopic constants of the bound states are determined and in good agreement with the available theoretical and experimental results. The permanent dipole moments (PDMs) of Λ-S states and the spin-orbit (SO) matrix elements between Λ-S states are also computed. The results show that the abrupt changes of the PDMs and SO matrix elements come into being for the reason of the avoided crossing between the states with the same symmetry. In addition, the non-adiabatic couplings matrix elements between Λ-S states are also evaluated. Finally, the spin-orbit couplings (SOCs) for the low-lying states are considered with Breit-Pauli operator. The SOC effect makes the 7 Λ-S states of the ZnH+ cation split into 12 Ω states (Ω = Λ + Sz, in which Sz is projection of the total electron spin S along the internuclear Z-axis). For the (3)0+ state, the two energy minima exhibit in the potential, which could be attributed to the formation of the new avoided crossing point. The transition dipole moments (TDMs), Franck-Condon factors, and the radiative lifetimes of the selected transitions (2)0+-X0+, (3)0+-X0+, (2)1-X0+ and (3)1-X0+ have been reported.
Accurate potential energy functions, non-adiabatic and spin-orbit couplings in the ZnH(+) system.
Liang, Guiying; Liu, Xiaoting; Zhang, Xiaomei; Xu, Haifeng; Yan, Bing
2016-03-01
A high-level ab initio calculation on the ZnH(+) cation has been carried out with the multi-reference configuration interaction method plus Davison correction (MRCI+Q). The scalar relativistic effect is included by using the Douglas-Kroll-Hess (DKH) method. The calculated potential energy curves (PECs) of the 7 Λ-S states are associated with the dissociation limits of Zn(+)((2)Sg)+H((2)Sg), Zn((1)Sg)+H(+)((1)Sg), and Zn(+)((2)Pu)+H((2)Sg), respectively (The Λ-S state is labeled as (2S+1)Λ, in which Λ is the quantum number for the projection along the internuclear axis of the total electronic orbital angular momentum and S is the total electron spin). The spectroscopic constants of the bound states are determined and in good agreement with the available theoretical and experimental results. The permanent dipole moments (PDMs) of Λ-S states and the spin-orbit (SO) matrix elements between Λ-S states are also computed. The results show that the abrupt changes of the PDMs and SO matrix elements come into being for the reason of the avoided crossing between the states with the same symmetry. In addition, the non-adiabatic couplings matrix elements between Λ-S states are also evaluated. Finally, the spin-orbit couplings (SOCs) for the low-lying states are considered with Breit-Pauli operator. The SOC effect makes the 7 Λ-S states of the ZnH(+) cation split into 12 Ω states (Ω=Λ+Sz, in which Sz is projection of the total electron spin S along the internuclear Z-axis). For the (3)0(+) state, the two energy minima exhibit in the potential, which could be attributed to the formation of the new avoided crossing point. The transition dipole moments (TDMs), Franck-Condon factors, and the radiative lifetimes of the selected transitions (2)0(+)-X0(+), (3)0(+)-X0(+), (2)1-X0(+) and (3)1-X0(+) have been reported. PMID:26637984
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
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.