Real-time precise orbit determination of LEO satellites using a single-frequency GPS receiver: Preliminary results of Chinese SJ-9A satellite

NASA Astrophysics Data System (ADS)

Sun, Xiucong; Han, Chao; Chen, Pei

2017-10-01

Spaceborne Global Positioning System (GPS) receivers are widely used for orbit determination of low-Earth-orbiting (LEO) satellites. With the improvement of measurement accuracy, single-frequency receivers are recently considered for low-cost small satellite missions. In this paper, a Schmidt-Kalman filter which processes single-frequency GPS measurements and broadcast ephemerides is proposed for real-time precise orbit determination of LEO satellites. The C/A code and L1 phase are linearly combined to eliminate the first-order ionospheric effects. Systematic errors due to ionospheric delay residual, group delay variation, phase center variation, and broadcast ephemeris errors, are lumped together into a noise term, which is modeled as a first-order Gauss-Markov process. In order to reduce computational complexity, the colored noise is considered rather than estimated in the orbit determination process. This ensures that the covariance matrix accurately represents the distribution of estimation errors without increasing the dimension of the state vector. The orbit determination algorithm is tested with actual flight data from the single-frequency GPS receiver onboard China's small satellite Shi Jian-9A (SJ-9A). Preliminary results using a 7-h data arc on October 25, 2012 show that the Schmidt-Kalman filter performs better than the standard Kalman filter in terms of accuracy.

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.

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 with Angle-only Data from the First Korean Optical Satellite Tracking System, OWL-Net

NASA Astrophysics Data System (ADS)

Choi, J.; Jo, J.

2016-09-01

The optical satellite tracking data obtained by the first Korean optical satellite tracking system, Optical Wide-field patrol - Network (OWL-Net), had been examined for precision orbit determination. During the test observation at Israel site, we have successfully observed a satellite with Laser Retro Reflector (LRR) to calibrate the angle-only metric data. The OWL observation system is using a chopper equipment to get dense observation data in one-shot over 100 points for the low Earth orbit objects. After several corrections, orbit determination process was done with validated metric data. The TLE with the same epoch of the end of the first arc was used for the initial orbital parameter. Orbit Determination Tool Kit (ODTK) was used for an analysis of a performance of orbit estimation using the angle-only measurements. We have been developing batch style orbit estimator.

Maneuver Recovery Analysis for the Magnetospheric Multiscale Mission

NASA Technical Reports Server (NTRS)

Gramling, Cheryl; Carpenter, Russell; Volle, Michael; Lee, Taesul; Long, Anne

2007-01-01

The use of spacecraft formations creates new and more demanding requirements for orbit determination accuracy. In addition to absolute navigation requirements, there are typically relative navigation requirements that are based on the size or shape of the formation. The difficulty in meeting these requirements is related to the relative dynamics of the spacecraft orbits and the frequency of the formation maintenance maneuvers. This paper examines the effects of bi-weekly formation maintenance maneuvers on the absolute and relative orbit determination accuracy for the four-spacecraft Magnetospheric Multiscale (MMS) formation. Results are presented from high fidelity simulations that include the effects of realistic orbit determination errors in the maneuver planning process. Solutions are determined using a high accuracy extended Kalman filter designed for onboard navigation. Three different solutions are examined, considering the effects of process noise and measurement rate on the solutions.

Testing of Selected Geopotential Models in Terms of GOCE Satellite Orbit Determination Using Simulated GPS Observations

NASA Astrophysics Data System (ADS)

Bobojc, Andrzej; Drozyner, Andrzej

2016-04-01

This work contains a comparative study of performance of twenty geopotential models in an orbit estimation process of the satellite of the Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) mission. For testing, among others, such models as JYY_GOCE02S, ITG-GOCE02, ULUX_CHAMP2013S, GOGRA02S, ITG-GRACE2010S, EIGEN-51C, EGM2008, EGM96, JGM3, OSU91a, OSU86F were adopted. A special software package, called the Orbital Computation System (OCS), based on the classical method of least squares was used. In the frame of OCS, initial satellite state vector components are corrected in an iterative process, using the given geopotential model and the models describing the remaining gravitational perturbations. An important part of the OCS package is the 8th order Cowell numerical integration procedure, which enables a satellite orbit computation. Different sets of pseudorange simulations along reference GOCE satellite orbital arcs were obtained using real orbits of the Global Positioning System (GPS) satellites. These sets were the basic observation data used in the adjustment. The centimeter-accuracy Precise Science Orbit (PSO) for the GOCE satellite provided by the European Space Agency (ESA) was adopted as the GOCE reference orbit. Comparing various variants of the orbital solutions, the relative accuracy of geopotential models in an orbital aspect is determined. Full geopotential models were used in the adjustment process. However, the solutions were also determined taking into account truncated geopotential models. In such case, an accuracy of the orbit estimated was slightly enhanced. The obtained solutions refer to the orbital arcs with the lengths of 90-minute and 1-day.

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.

Improved solution accuracy for TDRSS-based TOPEX/Poseidon orbit determination

NASA Technical Reports Server (NTRS)

Doll, C. E.; Mistretta, G. D.; Hart, R. C.; Oza, D. H.; Bolvin, D. T.; Cox, C. M.; Nemesure, M.; Niklewski, D. J.; Samii, M. V.

1994-01-01

Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using a batch-least-squares estimator available in the Goddard Trajectory Determination System (GTDS) and an extended Kalman filter estimation system to process Tracking and Data Relay Satellite (TDRS) System (TDRSS) measurements. GTDS is the operational orbit determination system used by the FDD in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. The extended Kalman filter was implemented in an orbit determination analysis prototype system, closely related to the Real-Time Orbit Determination System/Enhanced (RTOD/E) system. In addition, the Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generated an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the geodynamics (GEODYN) orbit determination system with laser ranging and Doppler Orbitography and Radiopositioning integrated by satellite (DORIS) tracking measurements. The TOPEX/Poseidon trajectories were estimated for November 7 through November 11, 1992, the timeframe under study. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch-least-squares solutions were assessed based on the solution residuals, while the sequential solutions were assessed based on primarily the estimated covariances. The batch-least-squares and sequential orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 2 meters for the batch-least-squares and less than 13 meters for the sequential estimation solutions. After the sequential estimation solutions were processed with a smoother algorithm, position differences with POD orbit solutions of less than 7 meters were obtained. The differences among the POD, GTDS, and filter/smoother solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.

Selected Gravity Models in Terms of the fit to the GOCE Kinematic Orbit in the Dynamic Orbit Determination Process

NASA Astrophysics Data System (ADS)

Bobojć, Andrzej; Drożyner, Andrzej; Rzepecka, Zofia

2017-04-01

The work includes the comparison of performance of selected geopotential models in the dynamic orbit estimation of the satellite of the Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) mission. This was realized by fitting estimated orbital arcs to the official centimeter-accuracy GOCE kinematic orbit which is provided by the European Space Agency. The Cartesian coordinates of kinematic orbit were treated as observations in the orbit estimation. The initial satellite state vector components were corrected in an iterative process with respect to the J2000.0 inertial reference frame using the given geopotential model, the models describing the remaining gravitational perturbations and the solar radiation pressure. Taking the obtained solutions into account, the RMS values of orbital residuals were computed. These residuals result from the difference between the determined orbit and the reference one - the GOCE kinematic orbit. The performance of selected gravity models was also determined using various orbital arc lengths. Additionally, the RMS fit values were obtained for some gravity models truncated at given degree and order of spherical harmonic coefficients. The advantage of using the kinematic orbit is its independence from any a priori dynamical models. For the research such GOCE-independent gravity models as HUST-Grace2016s, ITU_GRACE16, ITSG-Grace2014s, ITSG-Grace2014k, GGM05S, Tongji-GRACE01, ULUX_CHAMP2013S, ITG-GRACE2010S, EIGEN-51C, EIGEN5S, EGM2008 and EGM96 were adopted.

TOPEX/POSEIDON operational orbit determination results using global positioning satellites

NASA Technical Reports Server (NTRS)

Guinn, J.; Jee, J.; Wolff, P.; Lagattuta, F.; Drain, T.; Sierra, V.

1994-01-01

Results of operational orbit determination, performed as part of the TOPEX/POSEIDON (T/P) Global Positioning System (GPS) demonstration experiment, are presented in this article. Elements of this experiment include the GPS satellite constellation, the GPS demonstration receiver on board T/P, six ground GPS receivers, the GPS Data Handling Facility, and the GPS Data Processing Facility (GDPF). Carrier phase and P-code pseudorange measurements from up to 24 GPS satellites to the seven GPS receivers are processed simultaneously with the GDPF software MIRAGE to produce orbit solutions of T/P and the GPS satellites. Daily solutions yield subdecimeter radial accuracies compared to other GPS, LASER, and DORIS precision orbit solutions.

The binary Asteroid 22 Kalliope: Linus orbit determination on the basis of speckle interferometric observations

NASA Astrophysics Data System (ADS)

Sokova, I. A.; Sokov, E. N.; Roschina, E. A.; Rastegaev, D. A.; Kiselev, A. A.; Balega, Yu. Yu.; Gorshanov, D. L.; Malogolovets, E. V.; Dyachenko, V. V.; Maksimov, A. F.

2014-07-01

In this paper we present the orbital elements of Linus satellite of 22 Kalliope asteroid. Orbital element determination is based on the speckle interferometry data obtained with the 6-m BTA telescope operated by SAO RAS. We processed 9 accurate positions of Linus orbiting around the main component of 22 Kalliope between 10 and 16 December, 2011. In order to determine the orbital elements of the Linus we have applied the direct geometric method. The formal errors are about 5 mas. This accuracy makes it possible to study the variations of the Linus orbital elements influenced by different perturbations over the course of time. Estimates of six classical orbital elements, such as the semi-major axis of the Linus orbit a = 1109 ± 6 km, eccentricity e = 0.016 ± 0.004, inclination i = 101° ± 1° to the ecliptic plane and others, are presented in this work.

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.

Parameter estimation supplement to the Mission Analysis Evaluation and Space Trajectory Operations program (MAESTRO)

NASA Technical Reports Server (NTRS)

Bjorkman, W. S.; Uphoff, C. W.

1973-01-01

This Parameter Estimation Supplement describes the PEST computer program and gives instructions for its use in determination of lunar gravitation field coefficients. PEST was developed for use in the RAE-B lunar orbiting mission as a means of lunar field recovery. The observations processed by PEST are short-arc osculating orbital elements. These observations are the end product of an orbit determination process obtained with another program. PEST's end product it a set of harmonic coefficients to be used in long-term prediction of the lunar orbit. PEST employs some novel techniques in its estimation process, notably a square batch estimator and linear variational equations in the orbital elements (both osculating and mean) for measurement sensitivities. The program's capabilities are described, and operating instructions and input/output examples are given. PEST utilizes MAESTRO routines for its trajectory propagation. PEST's program structure and subroutines which are not common to MAESTRO are described. Some of the theoretical background information for the estimation process, and a derivation of linear variational equations for the Method 7 elements are included.

The possible effect of reaction wheel unloading on orbit determination for Chang'E-1 lunar mission

NASA Astrophysics Data System (ADS)

Jianguo, Yan; Jingsong, Ping; Fei, Li

During the flight of 3-axis stabilized lunar orbiter i e SELENE main orbiter Chang E-1 due to the overflow of the accumulated angular momentum the reaction-wheel will be unloaded during certain period so as to release the angular momentum for initialization Then the momentum wheel will be reloaded for satellite attitude measurement and control Above action will not only change the attitude but also change the orbit of the spacecraft Assuming the reaction-wheel unloading is carried out twice a day according to the current engineering designation and plan for SELENE main orbiter and Chang E-1 missions considering the algebra configuration of the tracking stations the Moon and the lunar orbiter the orbit determination is simulated for 14 days evolution of lunar orbiter In the simulation the satellite orbit is generated using GEODYNII code Based on the generated orbit the common view time period of the satellite by VLBI and USB network in every day is computed the orbit determination is processed for all the arcs of the orbit The orbit determination result of 28 orbits in 14 days is provided The orbits cover most of the possible geometrical configuration among orbiter the Moon and the tracking network The analysis here can benefit the tracking designation and plan for Chang E-1 mission

Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)

NASA Technical Reports Server (NTRS)

Mesarch, Michael A.; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James

2007-01-01

This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF's orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.

Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)

NASA Technical Reports Server (NTRS)

Mesarch, Michael; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James

2007-01-01

This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF s orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.

Ionospheric refraction effects on orbit determination using the orbit determination error analysis system

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.

Method of resolving radio phase ambiguity in satellite orbit determination

NASA Technical Reports Server (NTRS)

Councelman, Charles C., III; Abbot, Richard I.

1989-01-01

For satellite orbit determination, the most accurate observable available today is microwave radio phase, which can be differenced between observing stations and between satellites to cancel both transmitter- and receiver-related errors. For maximum accuracy, the integer cycle ambiguities of the doubly differenced observations must be resolved. To perform this ambiguity resolution, a bootstrapping strategy is proposed. This strategy requires the tracking stations to have a wide ranging progression of spacings. By conventional 'integrated Doppler' processing of the observations from the most widely spaced stations, the orbits are determined well enough to permit resolution of the ambiguities for the most closely spaced stations. The resolution of these ambiguities reduces the uncertainty of the orbit determination enough to enable ambiguity resolution for more widely spaced stations, which further reduces the orbital uncertainty. In a test of this strategy with six tracking stations, both the formal and the true errors of determining Global Positioning System satellite orbits were reduced by a factor of 2.

Detection of Unknown LEO Satellite Using Radar Measurements

NASA Astrophysics Data System (ADS)

Kamensky, S.; Samotokhin, A.; Khutorovsky, Z.; Alfriend, T.

While processing of the radar information aimed at satellite catalog maintenance some measurements do not correlate with cataloged and tracked satellites. These non-correlated measurements participate in the detection (primary orbit determination) of new (not cataloged) satellites. The satellite is considered newly detected when it is missing in the catalog and the primary orbit determination on the basis of the non-correlated measurements provides the accuracy sufficient for reliable correlation of future measurements. We will call this the detection condition. One non-correlated measurement in real conditions does not have enough accuracy and thus does not satisfy the detection condition. Two measurements separated by a revolution or more normally provides orbit determination with accuracy sufficient for selection of other measurements. However, it is not always possible to say with high probability (close to 1) that two measurements belong to one satellite. Three measurements for different revolutions, which are included into one orbit, have significantly higher chances to belong to one satellite. Thus the suggested detection (primary orbit determination) algorithm looks for three uncorrelated measurements in different revolutions for which we can determine the orbit inscribing them. The detection procedure based on search for the triplets is rather laborious. Thus only relatively high efficiency can be the reason for its practical implementation. The work presents the detailed description of the suggested detection procedure based on the search for triplets of uncorrelated measurements (for radar measurements). The break-ups of the tracked satellites provide the most difficult conditions for the operation of the detection algorithm and reveal explicitly its characteristics. The characteristics of time efficiency and reliability of the detected orbits are of maximum interest. Within this work we suggest to determine these characteristics using simulation of break-ups with further acquisition of measurements generated by the fragments. In particular, using simulation we can not only evaluate the characteristics of the algorithm but adjust its parameters for certain conditions: the orbit of the fragmented satellite, the features of the break-up, capabilities of detection radars etc. We describe the algorithm performing the simulation of radar measurements produced by the fragments of the parent satellite. This algorithm accounts of the basic factors affecting the characteristics of time efficiency and reliability of the detection. The catalog maintenance algorithm includes two major components detection and tracking. These are two processes permanently interacting with each other. This is actually in place for the processing of real radar data. The simulation must take this into account since one cannot obtain reliable characteristics of detection procedure simulating only this process. Thus we simulated both processes in their interaction. The work presents the results of simulation for the simplest case of a break-up in near-circular orbit with insignificant atmospheric drag. The simulations show rather high efficiency. We demonstrate as well that the characteristics of time efficiency and reliability of determined orbits essentially depend on the density of the observed break-up fragments.

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.

Analysis of Preferred Directions in Phase Space for Tidal Measurements at Europa

NASA Astrophysics Data System (ADS)

Boone, D.; Scheeres, D. J.

2012-12-01

The NASA Jupiter Europa Orbiter mission requires a circular, near-polar orbit to measure Europa's Love numbers, geophysical coefficients which give insight into whether a liquid ocean exists. This type of orbit about planetary satellites is known to be unstable. The effects of Jupiter's tidal gravity are seen in changes in Europa's gravity field and surface deformation, which are sensed through doppler tracking over time and altimetry measurements respectively. These two measurement types separately determine the h and k Love numbers, a combination of which bounds how thick the ice shell of Europa is and whether liquid water is present. This work shows how the properties of an unstable periodic orbit about Europa generate preferred measurement directions in position and velocity phase space for the orbit determination process. We generate an error covariance over seven days for the orbiter state and science parameters using a periodic orbit and then disperse the orbit initial conditions in a Monte Carlo simulation according to this covariance. The dispersed orbits are shown to have a bias toward longer lifetimes and we discuss this as an effect of the stable and unstable manifolds of the periodic orbit. Using an epoch formulation of a square-root information filter, measurements aligned with the unstable manifold mapped back in time add more information to the orbit determination process than measurements aligned with the stable manifold. This corresponds to a contraction in the uncertainty of the estimate of the desired parameters, including the Love numbers. We demonstrate this mapping mathematically using a representation of the State Transition Matrix involving its eigenvectors and eigenvalues. Then using the properties of left and right eigenvectors, we show how measurements in the orbit determination process are mapped in time leading to a concentration of information at epoch. We present examples of measurements taken on different time schedules to show the effect of preferred phase space directions in the estimation process. Manifold coordinate decomposition is applied to the orbit initial conditions as well as measurement partials in the filter to show the alignment of each with the stable and unstable manifolds of the periodic orbit. The connection between orbit lifetime and regions of increased information density in phase space is made using the properties of these manifolds. Low altitude, near-polar periodic orbits with these characteristics are discussed along with the estimation results for the Love numbers, orbiter state, and orbit lifetime. Different measurement schedules and the resulting estimation performance are presented along with an analysis of information content for single measurements with respect to manifold alignment. These results allow more sensitive estimation of the tidal Love numbers and may allow measurements to be taken less frequently or compensate for corrupted data arcs. Other measurement types will be mapped in the same way using the State Transition matrix and have increased information density at epoch if aligned with the unstable manifold. In the same way, these results are applicable to planetary satellite orbiters about Enceladus or Dione since they share the governing equations of motion and properties of unstable periodic orbits.

Error analysis of real time and post processed or bit determination of GFO using GPS tracking

NASA Technical Reports Server (NTRS)

Schreiner, William S.

1991-01-01

The goal of the Navy's GEOSAT Follow-On (GFO) mission is to map the topography of the world's oceans in both real time (operational) and post processed modes. Currently, the best candidate for supplying the required orbit accuracy is the Global Positioning System (GPS). The purpose of this fellowship was to determine the expected orbit accuracy for GFO in both the real time and post-processed modes when using GPS tracking. This report presents the work completed through the ending date of the fellowship.

Orbit determination of the Next-Generation Beidou satellites with Intersatellite link measurements and a priori orbit constraints

NASA Astrophysics Data System (ADS)

Ren, Xia; Yang, Yuanxi; Zhu, Jun; Xu, Tianhe

2017-11-01

Intersatellite Link (ISL) technology helps to realize the auto update of broadcast ephemeris and clock error parameters for Global Navigation Satellite System (GNSS). ISL constitutes an important approach with which to both improve the observation geometry and extend the tracking coverage of China's Beidou Navigation Satellite System (BDS). However, ISL-only orbit determination might lead to the constellation drift, rotation, and even lead to the divergence in orbit determination. Fortunately, predicted orbits with good precision can be used as a priori information with which to constrain the estimated satellite orbit parameters. Therefore, the precision of satellite autonomous orbit determination can be improved by consideration of a priori orbit information, and vice versa. However, the errors of rotation and translation in a priori orbit will remain in the ultimate result. This paper proposes a constrained precise orbit determination (POD) method for a sub-constellation of the new Beidou satellite constellation with only a few ISLs. The observation model of dual one-way measurements eliminating satellite clock errors is presented, and the orbit determination precision is analyzed with different data processing backgrounds. The conclusions are as follows. (1) With ISLs, the estimated parameters are strongly correlated, especially the positions and velocities of satellites. (2) The performance of determined BDS orbits will be improved by the constraints with more precise priori orbits. The POD precision is better than 45 m with a priori orbit constrain of 100 m precision (e.g., predicted orbits by telemetry tracking and control system), and is better than 6 m with precise priori orbit constraints of 10 m precision (e.g., predicted orbits by international GNSS monitoring & Assessment System (iGMAS)). (3) The POD precision is improved by additional ISLs. Constrained by a priori iGMAS orbits, the POD precision with two, three, and four ISLs is better than 6, 3, and 2 m, respectively. (4) The in-plane link and out-of-plane link have different contributions to observation configuration and system observability. The POD with weak observation configuration (e.g., one in-plane link and one out-of-plane link) should be tightly constrained with a priori orbits.

Accuracy of Satellite Optical Observations and Precise Orbit Determination

NASA Astrophysics Data System (ADS)

Shakun, L.; Koshkin, N.; Korobeynikova, E.; Strakhova, S.; Dragomiretsky, V.; Ryabov, A.; Melikyants, S.; Golubovskaya, T.; Terpan, S.

The monitoring of low-orbit space objects (LEO-objects) is performed in the Astronomical Observatory of Odessa I.I. Mechnikov National University (Ukraine) for many years. Decades-long archives of these observations are accessible within Ukrainian network of optical observers (UMOS). In this work, we give an example of orbit determination for the satellite with the 1500-km height of orbit based on angular observations in our observatory (Int. No. 086). For estimation of the measurement accuracy and accuracy of determination and propagation of satellite position, we analyze the observations of Ajisai satellite with the well-determined orbit. This allows making justified conclusions not only about random errors of separate measurements, but also to analyze the presence of systematic errors, including external ones to the measurement process. We have shown that the accuracy of one measurement has the standard deviation about 1 arcsec across the track and 1.4 arcsec along the track and systematical shifts in measurements of one track do not exceed 0.45 arcsec. Ajisai position in the interval of the orbit fitting is predicted with accuracy better than 30 m along the orbit and better than 10 m across the orbit for any its point.

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.

Genetic Algorithm for Initial Orbit Determination with Too Short Arc (Continued)

NASA Astrophysics Data System (ADS)

Li, Xin-ran; Wang, Xin

2017-04-01

When the genetic algorithm is used to solve the problem of too short-arc (TSA) orbit determination, due to the difference of computing process between the genetic algorithm and the classical method, the original method for outlier deletion is no longer applicable. In the genetic algorithm, the robust estimation is realized by introducing different loss functions for the fitness function, then the outlier problem of the TSA orbit determination is solved. Compared with the classical method, the genetic algorithm is greatly simplified by introducing in different loss functions. Through the comparison on the calculations of multiple loss functions, it is found that the least median square (LMS) estimation and least trimmed square (LTS) estimation can greatly improve the robustness of the TSA orbit determination, and have a high breakdown point.

Estimating maneuvers for precise relative orbit determination using GPS

NASA Astrophysics Data System (ADS)

Allende-Alba, Gerardo; Montenbruck, Oliver; Ardaens, Jean-Sébastien; Wermuth, Martin; Hugentobler, Urs

2017-01-01

Precise relative orbit determination is an essential element for the generation of science products from distributed instrumentation of formation flying satellites in low Earth orbit. According to the mission profile, the required formation is typically maintained and/or controlled by executing maneuvers. In order to generate consistent and precise orbit products, a strategy for maneuver handling is mandatory in order to avoid discontinuities or precision degradation before, after and during maneuver execution. Precise orbit determination offers the possibility of maneuver estimation in an adjustment of single-satellite trajectories using GPS measurements. However, a consistent formulation of a precise relative orbit determination scheme requires the implementation of a maneuver estimation strategy which can be used, in addition, to improve the precision of maneuver estimates by drawing upon the use of differential GPS measurements. The present study introduces a method for precise relative orbit determination based on a reduced-dynamic batch processing of differential GPS pseudorange and carrier phase measurements, which includes maneuver estimation as part of the relative orbit adjustment. The proposed method has been validated using flight data from space missions with different rates of maneuvering activity, including the GRACE, TanDEM-X and PRISMA missions. The results show the feasibility of obtaining precise relative orbits without degradation in the vicinity of maneuvers as well as improved maneuver estimates that can be used for better maneuver planning in flight dynamics operations.

Potential for on-orbit manufacture of large space structures using the pultrusion process

NASA Technical Reports Server (NTRS)

Wilson, Maywood L.; Macconochie, Ian O.; Johnson, Gary S.

1987-01-01

On-orbit manufacture of lightweight, high-strength, advanced-composite structures using the pultrusion process is proposed. This process is adaptable to a zero-gravity environment by using preimpregnated graphite-fiber reinforcement systems. The reinforcement material is preimpregnated with a high-performance thermoplastic resin at a ground station, is coiled on spools for compact storage, and is transported into Earth orbit. A pultrusion machine is installed in the Shuttle cargo bay from which very long lengths of the desired structure is fabricated on-orbit. Potential structural profiles include rods, angles, channels, hat sections, tubes, honeycomb-cored panels, and T, H, and I beams. A potential pultrudable thermoplastic/graphite composite material is presented as a model for determining the effect on Earth-to-orbit package density of an on-orbit manufacture, the package density is increased by 132 percent, and payload volume requirement is decreased by 56.3 percent. The fabrication method has the potential for on-orbit manufacture of structural members for space platforms, large space antennas, and long tethers.

Method and apparatus for relative navigation using reflected GPS signals

NASA Technical Reports Server (NTRS)

Cohen, Ian R. (Inventor); Boegner, Jr., Gregory J. (Inventor)

2010-01-01

A method and system to passively navigate an orbiting moving body towards an orbiting target using reflected GPS signals. A pair of antennas is employed to receive both direct signals from a plurality of GPS satellites and a second antenna to receive GPS signals reflected off an orbiting target. The direct and reflected signals are processed and compared to determine the relative distance and position of the orbiting moving body relative to the orbiting target.

THE STATISTICAL MECHANICS OF PLANET ORBITS

DOE Office of Scientific and Technical Information (OSTI.GOV)

Tremaine, Scott, E-mail: tremaine@ias.edu

2015-07-10

The final “giant-impact” phase of terrestrial planet formation is believed to begin with a large number of planetary “embryos” on nearly circular, coplanar orbits. Mutual gravitational interactions gradually excite their eccentricities until their orbits cross and they collide and merge; through this process the number of surviving bodies declines until the system contains a small number of planets on well-separated, stable orbits. In this paper we explore a simple statistical model for the orbit distribution of planets formed by this process, based on the sheared-sheet approximation and the ansatz that the planets explore uniformly all of the stable region ofmore » phase space. The model provides analytic predictions for the distribution of eccentricities and semimajor axis differences, correlations between orbital elements of nearby planets, and the complete N-planet distribution function, in terms of a single parameter, the “dynamical temperature,” that is determined by the planetary masses. The predicted properties are generally consistent with N-body simulations of the giant-impact phase and with the distribution of semimajor axis differences in the Kepler catalog of extrasolar planets. A similar model may apply to the orbits of giant planets if these orbits are determined mainly by dynamical evolution after the planets have formed and the gas disk has disappeared.« less

Improving Fermi Orbit Determination and Prediction in an Uncertain Atmospheric Drag Environment

NASA Technical Reports Server (NTRS)

Vavrina, Matthew A.; Newman, Clark P.; Slojkowski, Steven E.; Carpenter, J. Russell

2014-01-01

Orbit determination and prediction of the Fermi Gamma-ray Space Telescope trajectory is strongly impacted by the unpredictability and variability of atmospheric density and the spacecraft's ballistic coefficient. Operationally, Global Positioning System point solutions are processed with an extended Kalman filter for orbit determination, and predictions are generated for conjunction assessment with secondary objects. When these predictions are compared to Joint Space Operations Center radar-based solutions, the close approach distance between the two predictions can greatly differ ahead of the conjunction. This work explores strategies for improving prediction accuracy and helps to explain the prediction disparities. Namely, a tuning analysis is performed to determine atmospheric drag modeling and filter parameters that can improve orbit determination as well as prediction accuracy. A 45% improvement in three-day prediction accuracy is realized by tuning the ballistic coefficient and atmospheric density stochastic models, measurement frequency, and other modeling and filter parameters.

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.

Comparison of TOPEX/Poseidon orbit determination solutions obtained by the Goddard Space Flight Center Flight Dynamics Division and Precision Orbit Determination Teams

NASA Technical Reports Server (NTRS)

Doll, C.; Mistretta, G.; Hart, R.; Oza, D.; Cox, C.; Nemesure, M.; Bolvin, D.; Samii, Mina V.

1993-01-01

Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using the Goddard Trajectory Determination System (GTDS) and a real-time extended Kalman filter estimation system to process Tracking Data and Relay Satellite (TDRS) System (TDRSS) measurements in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. GTDS is the operational orbit determination system used by the FDD, and the extended Kalman fliter was implemented in an analysis prototype system, the Real-Time Orbit Determination System/Enhanced (RTOD/E). The Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generates an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the Geodynamics (GEODYN) orbit determination system with laser ranging tracking data. The TOPEX/Poseidon trajectories were estimated for the October 22 - November 1, 1992, timeframe, for which the latest preliminary POD results were available. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch cases were assessed using overlap comparisons, while the sequential cases were assessed with covariances and the first measurement residuals. The batch least-squares and forward-filtered RTOD/E orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 10 meters (m) for the batch least squares and less than 18 m for the sequential estimation solutions. The differences among the POD, GTDS, and RTOD/E solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.

Research on the impact factors of GRACE precise orbit determination by dynamic method

NASA Astrophysics Data System (ADS)

Guo, Nan-nan; Zhou, Xu-hua; Li, Kai; Wu, Bin

2018-07-01

With the successful use of GPS-only-based POD (precise orbit determination), more and more satellites carry onboard GPS receivers to support their orbit accuracy requirements. It provides continuous GPS observations in high precision, and becomes an indispensable way to obtain the orbit of LEO satellites. Precise orbit determination of LEO satellites plays an important role for the application of LEO satellites. Numerous factors should be considered in the POD processing. In this paper, several factors that impact precise orbit determination are analyzed, namely the satellite altitude, the time-variable earth's gravity field, the GPS satellite clock error and accelerometer observation. The GRACE satellites provide ideal platform to study the performance of factors for precise orbit determination using zero-difference GPS data. These factors are quantitatively analyzed on affecting the accuracy of dynamic orbit using GRACE observations from 2005 to 2011 by SHORDE software. The study indicates that: (1) with the altitude of the GRACE satellite is lowered from 480 km to 460 km in seven years, the 3D (three-dimension) position accuracy of GRACE satellite orbit is about 3˜4 cm based on long spans data; (2) the accelerometer data improves the 3D position accuracy of GRACE in about 1 cm; (3) the accuracy of zero-difference dynamic orbit is about 6 cm with the GPS satellite clock error products in 5 min sampling interval and can be raised to 4 cm, if the GPS satellite clock error products with 30 s sampling interval can be adopted. (4) the time-variable part of earth gravity field model improves the 3D position accuracy of GRACE in about 0.5˜1.5 cm. Based on this study, we quantitatively analyze the factors that affect precise orbit determination of LEO satellites. This study plays an important role to improve the accuracy of LEO satellites orbit determination.

An intelligent interface for satellite operations: Your Orbit Determination Assistant (YODA)

NASA Technical Reports Server (NTRS)

Schur, Anne

1988-01-01

An intelligent interface is often characterized by the ability to adapt evaluation criteria as the environment and user goals change. Some factors that impact these adaptations are redefinition of task goals and, hence, user requirements; time criticality; and system status. To implement adaptations affected by these factors, a new set of capabilities must be incorporated into the human-computer interface design. These capabilities include: (1) dynamic update and removal of control states based on user inputs, (2) generation and removal of logical dependencies as change occurs, (3) uniform and smooth interfacing to numerous processes, databases, and expert systems, and (4) unobtrusive on-line assistance to users of concepts were applied and incorporated into a human-computer interface using artificial intelligence techniques to create a prototype expert system, Your Orbit Determination Assistant (YODA). YODA is a smart interface that supports, in real teime, orbit analysts who must determine the location of a satellite during the station acquisition phase of a mission. Also described is the integration of four knowledge sources required to support the orbit determination assistant: orbital mechanics, spacecraft specifications, characteristics of the mission support software, and orbit analyst experience. This initial effort is continuing with expansion of YODA's capabilities, including evaluation of results of the orbit determination task.

HY-2A altimetry satellite GPS orbits processing and performances

NASA Astrophysics Data System (ADS)

Mercier, F.; Houry, S.; Couhert, A.; Cerri, L.

2012-04-01

The Chinese HY-2A altimetry satellite is on the mission orbit since 1st october 2011. This satellite uses a Doris receiver (French cooperation), a GPS receiver and a SLR retro-reflector for the precise orbit determination. The GPS is a dual frequency semi-codeless receiver. Precise orbits are computed at CNES on the basis of 7 days arcs since the beginning of the mission (repeat cycle is 14 days). This presentation describes the current processing performed at CNES for this satellite. The GPS only orbits perform very well and are compared with the Doris only orbits (floating ambiguity resolution, as for Jason 1 and 2). SLR measurements are also available at ILRS, and allow an external validation of the actual radial orbit performance. This talk adresses the current status of POE solutions and the prospects for improvement based on the preliminary analysis of the tracking data.

Evaluation and modeling of autonomous attitude thrust control for the Geostation Operational Environmental Satellite (GOES)-8 orbit determination

NASA Technical Reports Server (NTRS)

Forcey, W.; Minnie, C. R.; Defazio, R. L.

1995-01-01

The Geostationary Operational Environmental Satellite (GOES)-8 experienced a series of orbital perturbations from autonomous attitude control thrusting before perigee raising maneuvers. These perturbations influenced differential correction orbital state solutions determined by the Goddard Space Flight Center (GSFC) Goddard Trajectory Determination System (GTDS). The maneuvers induced significant variations in the converged state vector for solutions using increasingly longer tracking data spans. These solutions were used for planning perigee maneuvers as well as initial estimates for orbit solutions used to evaluate the effectiveness of the perigee raising maneuvers. This paper discusses models for the incorporation of attitude thrust effects into the orbit determination process. Results from definitive attitude solutions are modeled as impulsive thrusts in orbit determination solutions created for GOES-8 mission support. Due to the attitude orientation of GOES-8, analysis results are presented that attempt to absorb the effects of attitude thrusting by including a solution for the coefficient of reflectivity, C(R). Models to represent the attitude maneuvers are tested against orbit determination solutions generated during real-time support of the GOES-8 mission. The modeling techniques discussed in this investigation offer benefits to the remaining missions in the GOES NEXT series. Similar missions with large autonomous attitude control thrusting, such as the Solar and Heliospheric Observatory (SOHO) spacecraft and the INTELSAT series, may also benefit from these results.

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.

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.

Initial results of centralized autonomous orbit determination of the new-generation BDS satellites with inter-satellite link measurements

NASA Astrophysics Data System (ADS)

Tang, Chengpan; Hu, Xiaogong; Zhou, Shanshi; Liu, Li; Pan, Junyang; Chen, Liucheng; Guo, Rui; Zhu, Lingfeng; Hu, Guangming; Li, Xiaojie; He, Feng; Chang, Zhiqiao

2018-01-01

Autonomous orbit determination is the ability of navigation satellites to estimate the orbit parameters on-board using inter-satellite link (ISL) measurements. This study mainly focuses on data processing of the ISL measurements as a new measurement type and its application on the centralized autonomous orbit determination of the new-generation Beidou navigation satellite system satellites for the first time. The ISL measurements are dual one-way measurements that follow a time division multiple access (TDMA) structure. The ranging error of the ISL measurements is less than 0.25 ns. This paper proposes a derivation approach to the satellite clock offsets and the geometric distances from TDMA dual one-way measurements without a loss of accuracy. The derived clock offsets are used for time synchronization, and the derived geometry distances are used for autonomous orbit determination. The clock offsets from the ISL measurements are consistent with the L-band two-way satellite, and time-frequency transfer clock measurements and the detrended residuals vary within 0.5 ns. The centralized autonomous orbit determination is conducted in a batch mode on a ground-capable server for the feasibility study. Constant hardware delays are present in the geometric distances and become the largest source of error in the autonomous orbit determination. Therefore, the hardware delays are estimated simultaneously with the satellite orbits. To avoid uncertainties in the constellation orientation, a ground anchor station that "observes" the satellites with on-board ISL payloads is introduced into the orbit determination. The root-mean-square values of orbit determination residuals are within 10.0 cm, and the standard deviation of the estimated ISL hardware delays is within 0.2 ns. The accuracy of the autonomous orbits is evaluated by analysis of overlap comparison and the satellite laser ranging (SLR) residuals and is compared with the accuracy of the L-band orbits. The results indicate that the radial overlap differences between the autonomous orbits are less than 15.0 cm for the inclined geosynchronous orbit (IGSO) satellites and less than 10.0 cm for the MEO satellites. The SLR residuals are approximately 15.0 cm for the IGSO satellites and approximately 10.0 cm for the MEO satellites, representing an improvement over the L-band orbits.

A multi-satellite orbit determination problem in a parallel processing environment

NASA Technical Reports Server (NTRS)

Deakyne, M. S.; Anderle, R. J.

1988-01-01

The Engineering Orbit Analysis Unit at GE Valley Forge used an Intel Hypercube Parallel Processor to investigate the performance and gain experience of parallel processors with a multi-satellite orbit determination problem. A general study was selected in which major blocks of computation for the multi-satellite orbit computations were used as units to be assigned to the various processors on the Hypercube. Problems encountered or successes achieved in addressing the orbit determination problem would be more likely to be transferable to other parallel processors. The prime objective was to study the algorithm to allow processing of observations later in time than those employed in the state update. Expertise in ephemeris determination was exploited in addressing these problems and the facility used to bring a realism to the study which would highlight the problems which may not otherwise be anticipated. Secondary objectives were to gain experience of a non-trivial problem in a parallel processor environment, to explore the necessary interplay of serial and parallel sections of the algorithm in terms of timing studies, to explore the granularity (coarse vs. fine grain) to discover the granularity limit above which there would be a risk of starvation where the majority of nodes would be idle or under the limit where the overhead associated with splitting the problem may require more work and communication time than is useful.

Summary of Skylab S-193 altimeter altitude results. [orbit calculation and studies of the ocean bottom

NASA Technical Reports Server (NTRS)

Mcgoogan, J. T.; Leitao, C. D.; Wells, W. T.

1975-01-01

The SKYLAB S-193 altimeter altitude results are presented in a concise format for further use and analysis by the scientific community. The altimeter mission and instrumentation is described along with the altimeter processing techniques and values of parameters used for processing. The determination of reference orbits is discussed, and the tracking systems utilized are tabulated. Techniques for determining satellite pointing are presented and a tabulation of pointing for each data mission included. The geographical location, the ocean bottom topography, the altimeter-determined ocean surface topography, and the altimeter automatic gain control history is presented. Some typical applications of this data are suggested.

GEODYN operations description, volume 3. [computer program for estimation of orbit and geodetic parameters

NASA Technical Reports Server (NTRS)

Martin, T. V.; Mullins, N. E.

1972-01-01

The operating and set-up procedures for the multi-satellite, multi-arc GEODYN- Orbit Determination program are described. All system output is analyzed. The GEODYN Program is the nucleus of the entire GEODYN system. It is a definitive orbit and geodetic parameter estimation program capable of simultaneously processing observations from multiple arcs of multiple satellites. GEODYN has two modes of operation: (1) the data reduction mode and (2) the orbit generation mode.

Achieving and Validating the 1-centimeter Orbit: JASON-1 Precision Orbit Determination Using GPS, SLR, DORIS and Altimeter data

NASA Technical Reports Server (NTRS)

Luthcke, Scott B.; Zelensky, Nikita P.; Rowlands, David D.; Lemoine, Frank G.; Williams, Teresa A.

2003-01-01

Jason-1, launched on December 7, 2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. Jason-1 is no exception and has set a 1 cm radial orbit accuracy goal, which represents a factor of two improvement over what is currently being achieved for T/P. The challenge to precision orbit determination (POD) is both achieving the 1 cm radial orbit accuracy and evaluating and validating the performance of the 1 cm orbit. Fortunately, Jason-1 POD can rely on four independent tracking data types including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. In addition, to the enhanced GPS receiver, Jason-1 carries significantly improved SLR and DORIS tracking systems along with the altimeter itself. We demonstrate the 1 cm radial orbit accuracy goal has been achieved using GPS data alone in a reduced dynamic solution. It is also shown that adding SLR data to the GPS-based solutions improves the orbits even further. In order to assess the performance of these orbits it is necessary to process all of the available tracking data (GPS, SLR, DORIS and altimeter crossover differences) as either dependent or independent of the orbit solutions. It was also necessary to compute orbit solutions using various combinations of the four available tracking data in order to independently assess the orbit performance. Towards this end, we have greatly improved orbits determined solely from SLR+DORIS data by applying the reduced dynamic solution strategy. In addition, we have computed reduced dynamic orbits based on SLR, DORIS and crossover data that are a significant improvement over the SLR and DORIS based dynamic solutions. These solutions provide the best performing orbits for independent validation of the GPS-based reduced dynamic orbits.

Diagrams for comprehensive molecular orbital-based chemical reaction analyses: reactive orbital energy diagrams.

PubMed

Tsuneda, Takao; Singh, Raman Kumar; Chattaraj, Pratim Kumar

2018-05-15

Reactive orbital energy diagrams are presented as a tool for comprehensively performing orbital-based reaction analyses. The diagrams rest on the reactive orbital energy theory, which is the expansion of conceptual density functional theory (DFT) to an orbital energy-based theory. The orbital energies on the intrinsic reaction coordinates of fundamental reactions are calculated by long-range corrected DFT, which is confirmed to provide accurate orbital energies of small molecules, combining with a van der Waals (vdW) correlation functional, in order to examine the vdW effect on the orbital energies. By analysing the reactions based on the reactive orbital energy theory using these accurate orbital energies, it is found that vdW interactions significantly affect the orbital energies in the initial reaction processes and that more than 70% of reactions are determined to be initially driven by charge transfer, while the remaining structural deformation (dynamics)-driven reactions are classified into identity, cyclization and ring-opening, unimolecular dissociation, and H2 reactions. The reactive orbital energy diagrams, which are constructed using these results, reveal that reactions progress so as to delocalize the occupied reactive orbitals, which are determined as contributing orbitals and are usually not HOMOs, by hybridizing the unoccupied reactive orbitals, which are usually not LUMOs. These diagrams also raise questions about conventional orbital-based diagrams such as frontier molecular orbital diagrams, even for the well-established interpretation of Diels-Alder reactions.

Short-Arc Orbit Determination Results and Space Debris Test Observation of the OWL-Net

NASA Astrophysics Data System (ADS)

Choi, J.; Jo, J.; Yim, H.

Korea Astronomy and Space Science Institute had developed the Optical Wide-field patroL-Network (OWL-Net) for maintaining the domestic Low Earth Orbit satellites’ ephemeris and monitoring Geostationary Earth Orbit region. It also can be used to observe space debris. The orbit determination process was planned with batch least square orbit estimator for every week. The optical tracking window is very narrow, several times per week. Sequentialbatch type estimation strategy was attempted for more reliable orbit prediction. We compared the test operation results with Two Line Elements and CPF files to validate the system. This results can be used to estimate the performance of the OWL-Net operations. And also we had observation of the Astro-H debris. We got the dozens of photometric data of the Astro-H debris main part for a few seconds with the chopper system.

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.

Motion Parameters Determination of the SC and Phobos in the Project Phobos-Grunt

NASA Technical Reports Server (NTRS)

Akim, E. L.; Stepanyants, V. A.; Tuchin, A. G.; Shishov, V. A.

2007-01-01

The SC "Phobos-Grunt" flight is planned to 2009 in Russia with the purpose to deliver to the Earth the soil samples of the Mars satellite Phobos. The mission will pass under the following scheme [1-4]: the SC flight from the Earth to the Mars, the SC transit on the Mars satellite orbit, the motion round the Mars on the observation orbit and on the quasi-synchronous one [5], landing on Phobos, taking of a ground and start in the direction to the Earth. The implementation of complicated dynamical operations in the Phobos vicinity is foreseen by the project. The SC will be in a disturbance sphere of gravitational fields from the Sun, the Mars and the Phobos. The SC orbit determination is carried out on a totality of trajectory measurements executed from ground tracking stations and measurements of autonomous systems onboard space vehicle relatively the Phobos. As ground measurements the radio engineering measurements of range and range rate are used. There are possible as onboard optical observations of the Phobos by a television system and ranges from the SC up to the Phobos surface by laser locator. As soon as the Phobos orbit accuracy is insufficient for a solution of a problem of landing its orbit determination will be carried out together with determination of the SC orbit. Therefore the algorithms for joint improving of initial conditions of the SC and the Phobos are necessary to determine parameters of the SC relative the Phobos motion within a single dynamical motion model. After putting on the martial satellite orbit, on the Phobos observation orbit, on the quasi-synchronous orbit in the Phobos vicinity the equipment guidance and the following process of the SC orbit determination relatively Phobos requires a priori knowledge of the Phobos orbit parameters with sufficiently high precision. These parameters should be obtained beforehand using both all modern observations and historical ones.

Towards SI-traceable radio occultation excess phase processing with integrated uncertainty estimation for climate applications

NASA Astrophysics Data System (ADS)

Innerkofler, Josef; Pock, Christian; Kirchengast, Gottfried; Schwaerz, Marc; Jaeggi, Adrian; Schwarz, Jakob

2016-04-01

The GNSS Radio Occultation (RO) measurement technique is highly valuable for climate monitoring of the atmosphere as it provides accurate and precise measurements in the troposphere and stratosphere regions with global coverage, long-term stability, and virtually all-weather capability. The novel Reference Occultation Processing System (rOPS), currently under development at the WEGC at University of Graz aims to process raw RO measurements into essential climate variables, such as temperature, pressure, and tropospheric water vapor, in a way which is SI-traceable to the universal time standard and which includes rigorous uncertainty propagation. As part of this rOPS climate-quality processing system, accurate atmospheric excess phase profiles with new approaches integrating uncertainty propagation are derived from the raw occultation tracking data and orbit data. Regarding the latter, highly accurate orbit positions and velocities of the GNSS transmitter satellites and the RO receiver satellites in low Earth orbit (LEO) need to be determined, in order to enable high accuracy of the excess phase profiles. Using several representative test days of GPS orbit data from the CODE and IGS archives, which are available at accuracies of about 3 cm (position) / 0.03 mm/s (velocity), and employing Bernese 5.2 and Napeos 3.3.1 software packages for the LEO orbit determination of the CHAMP, GRACE, and MetOp RO satellites, we achieved robust SI-traced LEO orbit uncertainty estimates of about 5 cm (position) / 0.05 mm/s (velocity) for the daily orbits, including estimates of systematic uncertainty bounds and of propagated random uncertainties. For COSMIC RO satellites, we found decreased accuracy estimates near 10-15 cm (position) / 0.1-0.15 mm/s (velocity), since the characteristics of the small COSMIC satellite platforms and antennas provide somewhat less favorable orbit determination conditions. We present the setup of how we (I) used the Bernese and Napeos package in mutual cross-check for this purpose, (II) integrated satellite laser-ranging validation of the estimated systematic uncertainty bounds, (III) expanded the Bernese 5.2 software for propagating random uncertainties from the GPS orbit data and LEO navigation tracking data input to the LEO data output. Preliminary excess phase results including propagated uncertainty estimates will also be shown. Except for disturbed space weather conditions, we expect a robust performance at millimeter level for the derived excess phases, which after large-scale processing of the RO data of many years can provide a new SI-traced fundamental climate data record.

Precise orbit determination of BeiDou constellation based on BETS and MGEX network.

PubMed

Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu

2014-04-15

Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20 cm and 14 cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10 cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved.

Lunar gravity derived from long-period satellite motion, a proposed method

NASA Technical Reports Server (NTRS)

Ferrari, A. J.

1971-01-01

A method was devised to determine the spherical harmonic coefficients of the lunar gravity field. The method consists of a two-step data reduction and estimation process. Pseudo-Doppler data were generated simulating two different lunar orbits. The analysis included the perturbing effects of the L1 lunar gravity field, the earth, the sun, and solar radiation pressure. Orbit determinations were performed on these data and long-period orbital elements were obtained. The Kepler element rates from these solutions were used to recover L1 lunar gravity coefficients. Overall results of the experiment show that lunar gravity coefficients can be accurately determined and that the method is dynamically consistent with long-period perturbation theory.

Spacecraft Orbit Anomaly Representation Using Thrust-Fourier-Coefficients with Orbit Determination Toolbox

NASA Astrophysics Data System (ADS)

Ko, H.; Scheeres, D.

2014-09-01

Representing spacecraft orbit anomalies between two separate states is a challenging but an important problem in achieving space situational awareness for an active spacecraft. Incorporation of such a capability could play an essential role in analyzing satellite behaviors as well as trajectory estimation of the space object. A general way to deal with the anomaly problem is to add an estimated perturbing acceleration such as dynamic model compensation (DMC) into an orbit determination process based on pre- and post-anomaly tracking data. It is a time-consuming numerical process to find valid coefficients to compensate for unknown dynamics for the anomaly. Even if the orbit determination filter with DMC can crudely estimate an unknown acceleration, this approach does not consider any fundamental element of the unknown dynamics for a given anomaly. In this paper, a new way of representing a spacecraft anomaly using an interpolation technique with the Thrust-Fourier-Coefficients (TFCs) is introduced and several anomaly cases are studied using this interpolation method. It provides a very efficient way of reconstructing the fundamental elements of the dynamics for a given spacecraft anomaly. Any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver using an interpolation technique with the TFCs. Given unconnected orbit states between two epochs due to a spacecraft anomaly, it is possible to obtain a unique control law using the TFCs that is able to generate the desired secular behavior for the given orbital changes. This interpolation technique can capture the fundamental elements of combined unmodeled anomaly events. The interpolated orbit trajectory, using the TFCs compensating for a given anomaly, can be used to improve the quality of orbit fits through the anomaly period and therefore help to obtain a good orbit determination solution after the anomaly. Orbit Determination Toolbox (ODTBX) is modified to adapt this technique in order to verify the performance of this interpolation approach. Spacecraft anomaly cases are based on either single or multiple low or high thrust maneuvers and the unknown thrust accelerations are recovered and compared with the true thrust acceleration. The advantage of this approach is to easily append TFCs and its dynamics to the pre-built ODTBX, which enables us to blend post-anomaly tracking data to improve the performance of the interpolation representation in the absence of detailed information about a maneuver. It allows us to improve space situational awareness in the areas of uncertainty propagation, anomaly characterization and track correlation.

GNSS orbit determination by precise modeling of non-gravitational forces acting on satellite's body

NASA Astrophysics Data System (ADS)

Wielgosz, Agata; Kalarus, Maciej; Liwosz, Tomasz

2016-04-01

Satellites orbiting around Earth are affected by gravitational forces and non-gravitational perturbations (NGP). While the perturbations caused by gravitational forces, which are due to central body gravity (including high-precision geopotential field) and its changes (due to secular variations and tides), solar bodies attraction and relativistic effects are well-modeled, the perturbations caused by the non-gravitational forces are the most limiting factor in Precise Orbit Determination (POD). In this work we focused on very precise non-gravitational force modeling for medium Earth orbit satellites by applying the various models of solar radiation pressure including changes in solar irradiance and Earth/Moon shadow transition, Earth albedo and thermal radiation. For computing influence of aforementioned forces on spacecraft the analytical box-wing satellite model was applied. Smaller effects like antenna thrust or spacecraft thermal radiation were also included. In the process of orbit determination we compared the orbit with analytically computed NGP with the standard procedure in which CODE model is fitted for NGP recovery. We considered satellites from several systems and on different orbits and for different periods: when the satellite is all the time in full sunlight and when transits the umbra and penumbra regions.

A demonstration of high precision GPS orbit determination for geodetic applications

NASA Technical Reports Server (NTRS)

Lichten, S. M.; Border, J. S.

1987-01-01

High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.

The Copernicus POD Service and beyond: Scientific exploitation of the orbit-related data and products

NASA Astrophysics Data System (ADS)

Peter, Heike; Fernández, Jaime; Fernández, Carlos; Féménias, Pierre

2017-04-01

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. The orbital products are available through the dedicated Copernicus data hub. The Copernicus POD Service is supported by the Copernicus POD Quality Working Group (QWG) for the validation of the orbit product accuracy. The QWG is delivering independent orbit solutions for the satellites. The cross-comparison of all these orbit solutions is essential to monitor and to improve the orbit accuracy because for Sentinel-1 and -2 this is the only possibility to externally assess the quality of the orbits. Each of the Sentinel-1, -2, and -3 satellites carries dual-frequency GPS receivers delivering the necessary measurements for the precise orbit determination of the satellites. The Sentinel-3 satellites are additionally equipped with a DORIS (Doppler Orbitography and Radiopositioning Integrated by Satellite) receiver and a Laser Retro Reflector for Satellite Laser Ranging. These two additional observation techniques allow for independent validation of the GPS-derived orbit determination results and for studying biases between the different techniques. The scientific exploitation of the orbit determination and the corresponding input data is manifold. Sophisticated satellite macro models improve the modelling of the non-gravitational forces acting on the satellite. On the other hand, comparisons to orbits based on pure empirical modelling of the non-gravitational forces help to sort out deficiencies in the satellite geometry information. The dual-frequency GPS data delivered by the satellites can give valuable input for ionospheric studies important for Space Weather research. So-called kinematic orbits, being a time series of discrete satellite positions derived from GPS, may be used for the modelling of the time-variable low degree harmonics of the Earth's gravity field. This is very important to support filling the possible gap between the dedicated gravity field missions GRACE and GRACE Follow-on. Many other important research topics could be mentioned here as well. Therefore a broad scientific community could benefit of an open access not only to the operational orbits (which is partially available today), but also to the GPS observations, satellite attitude and other ancillary information to perform POD. This poster presents firstly the status of the Copernicus POD Service in terms of products generated, accuracy and timeliness of the operational orbital products and all potential inputs available. Then the main focus of the poster is to outline the possibilities for scientific exploitation of the orbit determination and the corresponding input data. The great scientific potential of these data is explained to confirm the need of making them publicly available for scientists.

Precise orbit determination for NASA's earth observing system using GPS (Global Positioning System)

NASA Technical Reports Server (NTRS)

Williams, B. G.

1988-01-01

An application of a precision orbit determination technique for NASA's Earth Observing System (EOS) using the Global Positioning System (GPS) is described. This technique allows the geometric information from measurements of GPS carrier phase and P-code pseudo-range to be exploited while minimizing requirements for precision dynamical modeling. The method combines geometric and dynamic information to determine the spacecraft trajectory; the weight on the dynamic information is controlled by adjusting fictitious spacecraft accelerations in three dimensions which are treated as first order exponentially time correlated stochastic processes. By varying the time correlation and uncertainty of the stochastic accelerations, the technique can range from purely geometric to purely dynamic. Performance estimates for this technique as applied to the orbit geometry planned for the EOS platforms indicate that decimeter accuracies for EOS orbit position may be obtainable. The sensitivity of the predicted orbit uncertainties to model errors for station locations, nongravitational platform accelerations, and Earth gravity is also presented.

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.

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.

A Comparison of JPDA and Belief Propagation for Data Association in SSA

NASA Astrophysics Data System (ADS)

Rutten, M.; Williams, J.; Gordon, N.; Jah, M.; Baldwin, J.; Stauch, J.

2014-09-01

The process of initial orbit determination, or catalogue maintenance, using a set of unlabeled observations requires a method of choosing which observation was due to which object. Realities of imperfect sensors mean that the association must be made in the presence of both missed detections and false alarms. Data association is not only essential to processing observations it can also be one of the most significant computational bottlenecks. The constrained admissible region multiple hypothesis filter (CAR-MHF) is an algorithm for initial orbit determination using short-arc observations of space objects. CAR-MHF has used joint probabilistic data association (JPDA), a well-established approach to multi-target data association. A recent development in the target tracking literature is the use of graphical models to formulate data association problems. Using an approximate inference algorithm, belief propagation (BP), on the graphical model results in an algorithm this is both computationally efficient and accurate. This paper compares CAR-MHF using JPDA and CAR-MHF using BP for the problem of initial orbit determination on a set of deep-space objects. The results of the analysis will show that by using the BP algorithm there are significant gains in computational load without any statistically significant loss in overall performance of the orbit determination.

Orbit Determination and Maneuver Detection Using Event Representation with Thrust-Fourier-Coefficients

NASA Astrophysics Data System (ADS)

Lubey, D.; Ko, H.; Scheeres, D.

The classical orbit determination (OD) method of dealing with unknown maneuvers is to restart the OD process with post-maneuver observations. However, it is also possible to continue the OD process through such unknown maneuvers by representing those unknown maneuvers with an appropriate event representation. It has been shown in previous work (Ko & Scheeres, JGCD 2014) that any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver connecting those two states using Thrust-Fourier-Coefficients (TFCs). Event representation using TFCs rigorously provides a unique control law that can generate the desired secular behavior for a given unknown maneuver. This paper presents applications of this representation approach to orbit prediction and maneuver detection problem across unknown maneuvers. The TFCs are appended to a sequential filter as an adjoint state to compensate unknown perturbing accelerations and the modified filter estimates the satellite state and thrust coefficients by processing OD across the time of an unknown maneuver. This modified sequential filter with TFCs is capable of fitting tracking data and maintaining an OD solution in the presence of unknown maneuvers. Also, the modified filter is found effective in detecting a sudden change in TFC values which indicates a maneuver. In order to illustrate that the event representation approach with TFCs is robust and sufficiently general to be easily adjustable, different types of measurement data are processed with the filter in a realistic LEO setting. Further, cases with mis-modeling of non-gravitational force are included in our study to verify the versatility and efficiency of our presented algorithm. Simulation results show that the modified sequential filter with TFCs can detect and estimate the orbit and thrust parameters in the presence of unknown maneuvers with or without measurement data during maneuvers. With no measurement data during maneuvers, the modified filter with TFCs uses an existing pre-maneuver orbit solution to compute a post-maneuver orbit solution by forcing TFCs to compensate for an unknown maneuver. With observation data available during maneuvers, maneuver start time and stop time is determined

Towards the GEOSAT Follow-On Precise Orbit Determination Goals of High Accuracy and Near-Real-Time Processing

NASA Technical Reports Server (NTRS)

Lemoine, Frank G.; Zelensky, Nikita P.; Chinn, Douglas S.; Beckley, Brian D.; Lillibridge, John L.

2006-01-01

The US Navy's GEOSAT Follow-On spacecraft (GFO) primary mission objective is to map the oceans using a radar altimeter. Satellite laser ranging data, especially in combination with altimeter crossover data, offer the only means of determining high-quality precise orbits. Two tuned gravity models, PGS7727 and PGS7777b, were created at NASA GSFC for GFO that reduce the predicted radial orbit through degree 70 to 13.7 and 10.0 mm. A macromodel was developed to model the nonconservative forces and the SLR spacecraft measurement offset was adjusted to remove a mean bias. Using these improved models, satellite-ranging data, altimeter crossover data, and Doppler data are used to compute both daily medium precision orbits with a latency of less than 24 hours. Final precise orbits are also computed using these tracking data and exported with a latency of three to four weeks to NOAA for use on the GFO Geophysical Data Records (GDR s). The estimated orbit precision of the daily orbits is between 10 and 20 cm, whereas the precise orbits have a precision of 5 cm.

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.

Extraction of stability and control derivatives from orbiter flight data

NASA Technical Reports Server (NTRS)

Iliff, Kenneth W.; Shafer, Mary F.

1993-01-01

The Space Shuttle Orbiter has provided unique and important information on aircraft flight dynamics. This information has provided the opportunity to assess the flight-derived stability and control derivatives for maneuvering flight in the hypersonic regime. In the case of the Space Shuttle Orbiter, these derivatives are required to determine if certain configuration placards (limitations on the flight envelope) can be modified. These placards were determined on the basis of preflight predictions and the associated uncertainties. As flight-determined derivatives are obtained, the placards are reassessed, and some of them are removed or modified. Extraction of the stability and control derivatives was justified by operational considerations and not by research considerations. Using flight results to update the predicted database of the orbiter is one of the most completely documented processes for a flight vehicle. This process followed from the requirement for analysis of flight data for control system updates and for expansion of the operational flight envelope. These results show significant changes in many important stability and control derivatives from the preflight database. This paper presents some of the stability and control derivative results obtained from Space Shuttle flights. Some of the limitations of this information are also examined.

Generalization of the Hartree-Fock approach to collision processes

NASA Astrophysics Data System (ADS)

Hahn, Yukap

1997-06-01

The conventional Hartree and Hartree-Fock approaches for bound states are generalized to treat atomic collision processes. All the single-particle orbitals, for both bound and scattering states, are determined simultaneously by requiring full self-consistency. This generalization is achieved by introducing two Ansäauttze: (a) the weak asymptotic boundary condition, which maintains the correct scattering energy and target orbitals with correct number of nodes, and (b) square integrable amputated scattering functions to generate self-consistent field (SCF) potentials for the target orbitals. The exact initial target and final-state asymptotic wave functions are not required and thus need not be specified a priori, as they are determined simultaneously by the SCF iterations. To check the asymptotic behavior of the solution, the theory is applied to elastic electron-hydrogen scattering at low energies. The solution is found to be stable and the weak asymptotic condition is sufficient to produce the correct scattering amplitudes. The SCF potential for the target orbital shows the strong penetration by the projectile electron during the collision, but the exchange term tends to restore the original form. Potential applicabilities of this extension are discussed, including the treatment of ionization and shake-off processes.

PSA: A program to streamline orbit determination for launch support operations

NASA Technical Reports Server (NTRS)

Legerton, V. N.; Mottinger, N. A.

1988-01-01

An interactive, menu driven computer program was written to streamline the orbit determination process during the critical launch support phase of a mission. Residing on a virtual memory minicomputer, this program retains the quantities in-core needed to obtain a least squares estimate of the spacecraft trajectory with interactive displays to assist in rapid radio metric data evaluation. Menu-driven displays allow real time filter and data strategy development. Graphical and tabular displays can be sent to a laser printer for analysis without exiting the program. Products generated by this program feed back to the main orbit determination program in order to further refine the estimate of the trajectory. The final estimate provides a spacecraft ephemeris which is transmitted to the mission control center and used for antenna pointing and frequency predict generation by the Deep Space Network. The development and implementation process of this program differs from that used for most other navigation software by allowing the users to check important operating features during development and have changes made as needed.

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.

Application of Numerical Integration and Data Fusion in Unit Vector Method

NASA Astrophysics Data System (ADS)

Zhang, J.

2012-01-01

The Unit Vector Method (UVM) is a series of orbit determination methods which are designed by Purple Mountain Observatory (PMO) and have been applied extensively. It gets the conditional equations for different kinds of data by projecting the basic equation to different unit vectors, and it suits for weighted process for different kinds of data. The high-precision data can play a major role in orbit determination, and accuracy of orbit determination is improved obviously. The improved UVM (PUVM2) promoted the UVM from initial orbit determination to orbit improvement, and unified the initial orbit determination and orbit improvement dynamically. The precision and efficiency are improved further. In this thesis, further research work has been done based on the UVM: Firstly, for the improvement of methods and techniques for observation, the types and decision of the observational data are improved substantially, it is also asked to improve the decision of orbit determination. The analytical perturbation can not meet the requirement. So, the numerical integration for calculating the perturbation has been introduced into the UVM. The accuracy of dynamical model suits for the accuracy of the real data, and the condition equations of UVM are modified accordingly. The accuracy of orbit determination is improved further. Secondly, data fusion method has been introduced into the UVM. The convergence mechanism and the defect of weighted strategy have been made clear in original UVM. The problem has been solved in this method, the calculation of approximate state transition matrix is simplified and the weighted strategy has been improved for the data with different dimension and different precision. Results of orbit determination of simulation and real data show that the work of this thesis is effective: (1) After the numerical integration has been introduced into the UVM, the accuracy of orbit determination is improved obviously, and it suits for the high-accuracy data of available observation apparatus. Compare with the classical differential improvement with the numerical integration, its calculation speed is also improved obviously. (2) After data fusion method has been introduced into the UVM, weighted distribution accords rationally with the accuracy of different kinds of data, all data are fully used and the new method is also good at numerical stability and rational weighted distribution.

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 a few centimeters with respect to high-quality JPL reference orbits. This shows a slightly better consistency compared to the case when using force models. A purely dynamic orbit, without estimating empirical accelerations thus only adjusting six state parameters and the bias and scale factors, gives an orbit fit for the GRACE B test case below the decimeter level. The in orbit calibrated accelerometer observations can be used to validate the modelled accelerations and estimated empirical accelerations computed with the GHOST tools. In along track direction they show the best resemblance, with a mean correlation coefficient of 93% for the same period. In radial and normal direction the correlation is smaller. During days of high solar activity the benefit of using accelerometer observations is clearly visible. The observations during these days show fluctuations which the modelled and empirical accelerations can not follow.

Demonstration of the DSST State Transition Matrix Time-Update Properties Using the Linux GTDS Program

DTIC Science & Technology

2011-09-01

by a single mean equinoctial element set . EGP Orbit Determination Test Cases Rev 25 14 All of the EGP test cases employ the same observation...the non-singular equinoctial mean elements is more linear and this has positive implications for orbit determination processes based on the semi...by a single mean equinoctial element set . 5. CONCLUSIONS The GTDS Semi-analytical Satellite Theory (DSST) architecture has been extended to

Investigation for improving Global Positioning System (GPS) orbits using a discrete sequential estimator and stochastic models of selected physical processes

NASA Technical Reports Server (NTRS)

Goad, Clyde C.; Chadwell, C. David

1993-01-01

GEODYNII is a conventional batch least-squares differential corrector computer program with deterministic models of the physical environment. Conventional algorithms were used to process differenced phase and pseudorange data to determine eight-day Global Positioning system (GPS) orbits with several meter accuracy. However, random physical processes drive the errors whose magnitudes prevent improving the GPS orbit accuracy. To improve the orbit accuracy, these random processes should be modeled stochastically. The conventional batch least-squares algorithm cannot accommodate stochastic models, only a stochastic estimation algorithm is suitable, such as a sequential filter/smoother. Also, GEODYNII cannot currently model the correlation among data values. Differenced pseudorange, and especially differenced phase, are precise data types that can be used to improve the GPS orbit precision. To overcome these limitations and improve the accuracy of GPS orbits computed using GEODYNII, we proposed to develop a sequential stochastic filter/smoother processor by using GEODYNII as a type of trajectory preprocessor. Our proposed processor is now completed. It contains a correlated double difference range processing capability, first order Gauss Markov models for the solar radiation pressure scale coefficient and y-bias acceleration, and a random walk model for the tropospheric refraction correction. The development approach was to interface the standard GEODYNII output files (measurement partials and variationals) with software modules containing the stochastic estimator, the stochastic models, and a double differenced phase range processing routine. Thus, no modifications to the original GEODYNII software were required. A schematic of the development is shown. The observational data are edited in the preprocessor and the data are passed to GEODYNII as one of its standard data types. A reference orbit is determined using GEODYNII as a batch least-squares processor and the GEODYNII measurement partial (FTN90) and variational (FTN80, V-matrix) files are generated. These two files along with a control statement file and a satellite identification and mass file are passed to the filter/smoother to estimate time-varying parameter states at each epoch, improved satellite initial elements, and improved estimates of constant parameters.

The orbits of the uranian satellites and rings, the gravity field of the uranian system, and the orientation of the pole of Uranus

DOE Office of Scientific and Technical Information (OSTI.GOV)

Jacobson, R. A., E-mail: robert.jacobson@jpl.nasa.gov

2014-11-01

French et al. determined the orbits of the Uranian rings, the orientation of the pole of Uranus, and the gravity harmonics of Uranus from Earth-based and Voyager ring occultations. Jacobson et al. determined the orbits of the Uranian satellites and the masses of Uranus and its satellites from Earth-based astrometry and observations acquired with the Voyager 2 spacecraft; they used the gravity harmonics and pole from French et al. Jacobson and Rush reconstructed the Voyager 2 trajectory and redetermined the Uranian system gravity parameters, satellite orbits, and ring orbits in a combined analysis of the data used previously augmented withmore » additional Earth-based astrometry. Here we report on an extension of that work that incorporates additional astrometry and ring occultations together with improved data processing techniques.« less

Classification of Stellar Orbits Near Corotation

NASA Astrophysics Data System (ADS)

Breet, Jessica; Daniel, Kathryne J.; Bryn Mawr College Galaxy Lab

2018-01-01

The process of radial migration is frequently invoked as an important process to spiral galaxy evolution, but the physical properties that determine the efficiency of radial migration are poorly defined. In order for a star to migrate radially it must first be trapped in a resonant orbit at the corotation radius of a spiral pattern. Stars in such trapped orbits have changing average orbital radii — and thus orbital angular momenta — without any change in orbital eccentricity. It follows that transient spiral patterns can permanently rearrange the distribution of orbital angular momentum in the disk without kinematically heating it. It is also known that orbits can also have a significant dynamical response at Lindblad Resonances (LRs), where the Ultraharmonic Lindblad Resonances (ULRs) have a lesser impact on the disk. The goal of our project is to examine and constrain the efficiency of radial migration via an investigation into whether or not stars in trapped orbits have a dynamical response at the ULRs. We produced a dataset of nearly 105 orbits with initial conditions across a range of radii and 2-D velocities. We then classified these orbits into four categories based on analytic criteria for whether or not they are in trapped orbits and/or cross the ULR over 1 gigayear. Preliminary investigations show that trapped orbits that also meet the ULR have a chaotic response, putting a potential limit on the efficiency of radial migration.

Flight dynamics facility operational orbit determination support for the ocean topography experiment

NASA Technical Reports Server (NTRS)

Bolvin, D. T.; Schanzle, A. F.; Samii, M. V.; Doll, C. E.

1991-01-01

The Ocean Topography Experiment (TOPEX/POSEIDON) mission is designed to determine the topography of the Earth's sea surface across a 3 yr period, beginning with launch in June 1992. The Goddard Space Flight Center Dynamics Facility has the capability to operationally receive and process Tracking and Data Relay Satellite System (TDRSS) tracking data. Because these data will be used to support orbit determination (OD) aspects of the TOPEX mission, the Dynamics Facility was designated to perform TOPEX operational OD. The scientific data require stringent OD accuracy in navigating the TOPEX spacecraft. The OD accuracy requirements fall into two categories: (1) on orbit free flight; and (2) maneuver. The maneuver OD accuracy requirements are of two types; premaneuver planning and postmaneuver evaluation. Analysis using the Orbit Determination Error Analysis System (ODEAS) covariance software has shown that, during the first postlaunch mission phase of the TOPEX mission, some postmaneuver evaluation OD accuracy requirements cannot be met. ODEAS results also show that the most difficult requirements to meet are those that determine the change in the components of velocity for postmaneuver evaluation.

Improved treatment of global positioning system force parameters in precise orbit determination applications

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.

Navigation study for low-altitude Earth satellites

NASA Technical Reports Server (NTRS)

Pastor, P. R.; Fang, B. T.; Yee, C. P.

1985-01-01

This document describes several navigation studies for low-altitude Earth satellites. The use of Global Positioning System Navigation Package data for LANDSAT-5 orbit determination is evaluated. In addition, a navigation analysis for the proposed Tracking and Data Aquisition System is presented. This analysis, based on simulations employing one-way Doppler data, is used to determine the agreement between the Research and Development Goddard Trajectory Determination System and the Sequential Error Analysis Program results. Properties of several geopotential error models are studied and an exploratory study of orbit smoother process noise is presented.

Precise orbit determination of BeiDou constellation based on BETS and MGEX network

PubMed Central

Lou, Yidong; Liu, Yang; Shi, Chuang; Yao, Xiuguang; Zheng, Fu

2014-01-01

Chinese BeiDou Navigation Satellite System is officially operational as a regional constellation with five Geostationary Earth Orbit (GEO) satellites, five Inclined Geosynchronous Satellite Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Observations from the BeiDou Experimental Tracking Stations (BETS) and the IGS Multi-GNSS Experiment (MGEX) network from 1 January to 31 March 2013 are processed for orbit determination of the BeiDou constellation. Various arc lengths and solar radiation pressure parameters are investigated. The reduced set of ECOM five-parameter model produces better performance than the full set of ECOM nine-parameter model for BeiDou IGSO and MEO. The orbit overlap for the middle days of 3-day arc solutions is better than 20 cm and 14 cm for IGSO and MEO in RMS, respectively. Satellite laser ranging residuals are better than 10 cm for both IGSO and MEO. For BeiDou GEO, the orbit overlap of several meters and satellite laser ranging residuals of several decimetres can be achieved. PMID:24733025

GEOS 3 data processing for the recovery of geoid undulations and gravity anomalies

NASA Technical Reports Server (NTRS)

Rapp, R. H.

1979-01-01

The paper discusses the analysis of GEOS 3 altimeter data for the determination of geoid heights and point and mean gravity anomalies. Methods are presented for determining the mean anomalies and mean undulations from the GEOS 3 altimeter data available by the end of September 1977 without having a complete set of precise orbits. The editing of the data is extensive to remove questionable data, although no filtering of the data is carried out. An adjustment process is carried out to eliminate orbit error and altimeter bias. Representative point anomaly values are computed to investigate anomaly behavior across the Bonin Trench and over the Patton seamounts.

FAST TRACK COMMUNICATION: Attosecond correlation dynamics during electron tunnelling from molecules

NASA Astrophysics Data System (ADS)

Walters, Zachary B.; Smirnova, Olga

2010-08-01

In this communication, we present an analytical theory of strong-field ionization of molecules, which takes into account the rearrangement of multiple interacting electrons during the ionization process. We show that such rearrangement offers an alternative pathway to the ionization of orbitals more deeply bound than the highest occupied molecular orbital. This pathway is not subject to the full exponential suppression characteristic of direct tunnel ionization from the deeper orbitals. The departing electron produces an 'attosecond correlation pulse' which controls the rearrangement during the tunnelling process. The shape and duration of this pulse are determined by the electronic structure of the relevant states, molecular orientation and laser parameters.

Global GNSS processing based on the raw observation approach

NASA Astrophysics Data System (ADS)

Strasser, Sebastian; Zehentner, Norbert; Mayer-Gürr, Torsten

2017-04-01

Many global navigation satellite system (GNSS) applications, e.g. Precise Point Positioning (PPP), require high-quality GNSS products, such as precise GNSS satellite orbits and clocks. These products are routinely determined by analysis centers of the International GNSS Service (IGS). The current processing methods of the analysis centers make use of the ionosphere-free linear combination to reduce the ionospheric influence. Some of the analysis centers also form observation differences, in general double-differences, to eliminate several additional error sources. The raw observation approach is a new GNSS processing approach that was developed at Graz University of Technology for kinematic orbit determination of low Earth orbit (LEO) satellites and subsequently adapted to global GNSS processing in general. This new approach offers some benefits compared to well-established approaches, such as a straightforward incorporation of new observables due to the avoidance of observation differences and linear combinations. This becomes especially important in view of the changing GNSS landscape with two new systems, the European system Galileo and the Chinese system BeiDou, currently in deployment. GNSS products generated at Graz University of Technology using the raw observation approach currently comprise precise GNSS satellite orbits and clocks, station positions and clocks, code and phase biases, and Earth rotation parameters. To evaluate the new approach, products generated using the Global Positioning System (GPS) constellation and observations from the global IGS station network are compared to those of the IGS analysis centers. The comparisons show that the products generated at Graz University of Technology are on a similar level of quality to the products determined by the IGS analysis centers. This confirms that the raw observation approach is applicable to global GNSS processing. Some areas requiring further work have been identified, enabling future improvements of the method.

A novel orbiter mission concept for venus with the EnVision proposal

NASA Astrophysics Data System (ADS)

de Oliveira, Marta R. R.; Gil, Paulo J. S.; Ghail, Richard

2018-07-01

In space exploration, planetary orbiter missions are essential to gain insight into planets as a whole, and to help uncover unanswered scientific questions. In particular, the planets closest to the Earth have been a privileged target of the world's leading space agencies. EnVision is a mission proposal designed for Venus and competing for ESA's next launch opportunity with the objective of studying Earth's closest neighbor. The main goal is to study geological and atmospheric processes, namely surface processes, interior dynamics and atmosphere, to determine the reasons behind Venus and Earth's radically different evolution despite the planets' similarities. To achieve these goals, the operational orbit selection is a fundamental element of the mission design process. The design of an orbit around Venus faces specific challenges, such as the impossibility of choosing Sun-synchronous orbits. In this paper, an innovative genetic algorithm optimization was applied to select the optimal orbit based on the parameters with more influence in the mission planning, in particular the mission duration and the coverage of sites of interest on the Venusian surface. The solution obtained is a near-polar circular orbit with an altitude of 259 km that enables the coverage of all priority targets almost two times faster than with the parameters considered before this study.

Constraints on the Efficiency of Radial Migration in Spiral Galaxies

NASA Astrophysics Data System (ADS)

Daniel, Kathryne J.; Wyse, Rosemary F. G.

2015-01-01

A transient spiral arm can permanently rearrange the orbital angular momentum of the stellar disk without inducing kinematic heating. This phenomenon is called radial migration because a star's orbital angular momentum determines its mean orbital radius. Should radial migration be an efficient process it could cause a large fraction of disk stars to experience significant changes in their individual orbital angular momenta on dynamically short timescales. Such scenarios have strong implications for the chemical, structural and kinematic evolution of disk galaxies. We have undertaken an investigation into the physical dependencies of the efficiency of radial migration on stellar kinematics and spiral structure. In order for a disk star to migrate radially, it must first be 'trapped' in a particular family of orbits, called horseshoe orbits, that occur near the radius of corotation with a spiral pattern. Thus far, the only analytic criterion for horseshoe orbits has been for stars with zero random orbital energy. We present our analytically derived 'capture criterion' for stars with some finite random orbital energy in a disk with a given rotation curve. Our capture criterion predict that trapping in a horseshoe orbit is primarily determined by whether or not the position of a star's mean orbital radius (determined by its orbital angular momentum) is within the 'capture region', the location and shape of which can be derived from the capture criterion. We visualize and confirm this prediction via numerically integrated orbits. We then apply our capture criterion to snap shot models of disk galaxies to determine (1) the radial distribution of the fraction of stars initially trapped in horseshoe orbits, and (2) the dependence of the total fraction of captured stars in the disk on the radial component of the stellar velocity dispersion (σR) and the amplitude of the spiral perturbation to the underlying potential at corotation. We here present a model of an exponential disk with a flat rotation curve where the initial fraction of stars trapped in horseshoe orbits falls with increasing velocity dispersion as exp[-σR^2].

A Deep Space Orbit Determination Software: Overview and Event Prediction Capability

NASA Astrophysics Data System (ADS)

Kim, Youngkwang; Park, Sang-Young; Lee, Eunji; Kim, Minsik

2017-06-01

This paper presents an overview of deep space orbit determination software (DSODS), as well as validation and verification results on its event prediction capabilities. DSODS was developed in the MATLAB object-oriented programming environment to support the Korea Pathfinder Lunar Orbiter (KPLO) mission. DSODS has three major capabilities: celestial event prediction for spacecraft, orbit determination with deep space network (DSN) tracking data, and DSN tracking data simulation. To achieve its functionality requirements, DSODS consists of four modules: orbit propagation (OP), event prediction (EP), data simulation (DS), and orbit determination (OD) modules. This paper explains the highest-level data flows between modules in event prediction, orbit determination, and tracking data simulation processes. Furthermore, to address the event prediction capability of DSODS, this paper introduces OP and EP modules. The role of the OP module is to handle time and coordinate system conversions, to propagate spacecraft trajectories, and to handle the ephemerides of spacecraft and celestial bodies. Currently, the OP module utilizes the General Mission Analysis Tool (GMAT) as a third-party software component for highfidelity deep space propagation, as well as time and coordinate system conversions. The role of the EP module is to predict celestial events, including eclipses, and ground station visibilities, and this paper presents the functionality requirements of the EP module. The validation and verification results show that, for most cases, event prediction errors were less than 10 millisec when compared with flight proven mission analysis tools such as GMAT and Systems Tool Kit (STK). Thus, we conclude that DSODS is capable of predicting events for the KPLO in real mission applications.

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 guidelines for determining optimal filter tuning parameters in a given sequential OD scenario for both covariance analysis and actual OD. Comparisons are also made with corresponding definitive OD results available from the TONS-EUVE analysis.

GOCE Precise Science Orbits for the Entire Mission and their Use for Gravity Field Recovery

NASA Astrophysics Data System (ADS)

Jäggi, Adrian; Bock, Heike; Meyer, Ulrich; Weigelt, Matthias

The Gravity field and steady-state Ocean Circulation Explorer (GOCE), ESA's first Earth Explorer Core Mission, was launched on March 17, 2009 into a sun-synchronous dusk-dawn orbit and re-entered into the Earth's atmosphere on November 11, 2013. It was equipped with a three-axis gravity gradiometer for high-resolution recovery of the Earth's gravity field, as well as with a 12-channel, dual-frequency Global Positioning System (GPS) receiver for precise orbit determination (POD), instrument time-tagging, and the determination of the long wavelength part of the Earth’s gravity field. A precise science orbit (PSO) product was provided during the entire mission by the GOCE High-level Processing Facility (HPF) from the GPS high-low Satellite-to-Satellite Tracking (hl-SST) data. We present the reduced-dynamic and kinematic PSO results for the entire mission period. Orbit comparisons and validations with independent Satellite Laser Ranging (SLR) measurements demonstrate the high quality of both orbit products being close to 2 cm 1-D RMS, but also reveal a correlation between solar activity, GPS data availability, and the quality of the orbits. We use the 1-sec kinematic positions of the GOCE PSO product for gravity field determination and present GPS-only solutions covering the entire mission period. The generated gravity field solutions reveal severe systematic errors centered along the geomagnetic equator, which may be traced back to the GPS carrier phase observations used for the kinematic orbit determination. The nature of the systematic errors is further investigated and reprocessed orbits free of systematic errors along the geomagnetic equator are derived. Eventually, the potential of recovering time variable signals from GOCE kinematic positions is assessed.

Orbit determination of highly elliptical Earth orbiters using improved Doppler data-processing modes

NASA Technical Reports Server (NTRS)

Estefan, J. A.

1995-01-01

A navigation error covariance analysis of four highly elliptical Earth orbits is described, with apogee heights ranging from 20,000 to 76,800 km and perigee heights ranging from 1,000 to 5,000 km. This analysis differs from earlier studies in that improved navigation data-processing modes were used to reduce the radio metric data. For this study, X-band (8.4-GHz) Doppler data were assumed to be acquired from two Deep Space Network radio antennas and reconstructed orbit errors propagated over a single day. Doppler measurements were formulated as total-count phase measurements and compared to the traditional formulation of differenced-count frequency measurements. In addition, an enhanced data-filtering strategy was used, which treated the principal ground system calibration errors affecting the data as filter parameters. Results suggest that a 40- to 60-percent accuracy improvement may be achievable over traditional data-processing modes in reconstructed orbit errors, with a substantial reduction in reconstructed velocity errors at perigee. Historically, this has been a regime in which stringent navigation requirements have been difficult to meet by conventional methods.

Monte Carlo Analysis as a Trajectory Design Driver for the TESS Mission

NASA Technical Reports Server (NTRS)

Nickel, Craig; Lebois, Ryan; Lutz, Stephen; Dichmann, Donald; Parker, Joel

2016-01-01

The Transiting Exoplanet Survey Satellite (TESS) will be injected into a highly eccentric Earth orbit and fly 3.5 phasing loops followed by a lunar flyby to enter a mission orbit with lunar 2:1 resonance. Through the phasing loops and mission orbit, the trajectory is significantly affected by lunar and solar gravity. We have developed a trajectory design to achieve the mission orbit and meet mission constraints, including eclipse avoidance and a 30-year geostationary orbit avoidance requirement. A parallelized Monte Carlo simulation was performed to validate the trajectory after injecting common perturbations, including launch dispersions, orbit determination errors, and maneuver execution errors. The Monte Carlo analysis helped identify mission risks and is used in the trajectory selection process.

NASA Advanced Concepts Office, Earth-To-Orbit Team Design Process and Tools

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Garcia, Jessica; Beers, Benjamin; Philips, Alan; Holt, James B.; Threet, Grady E., Jr.

2013-01-01

The Earth to Orbit (ETO) Team of the Advanced Concepts Office (ACO) at NASA Marshal Space Flight Center (MSFC) is considered the preeminent group to go to for prephase A and phase A concept definition. The ACO team has been at the forefront of a multitude of launch vehicle studies determining the future direction of the Agency as a whole due, in part, to their rapid turnaround time in analyzing concepts and their ability to cover broad trade spaces of vehicles in that limited timeframe. Each completed vehicle concept includes a full mass breakdown of each vehicle to tertiary subsystem components, along with a vehicle trajectory analysis to determine optimized payload delivery to specified orbital parameters, flight environments, and delta v capability. Additionally, a structural analysis of the vehicle based on material properties and geometries is performed as well as an analysis to determine the flight loads based on the trajectory outputs. As mentioned, the ACO Earth to Orbit Team prides themselves on their rapid turnaround time and often need to fulfill customer requests within limited schedule or little advanced notice. Due to working in this fast paced environment, the ETO team has developed some finely honed skills and methods to maximize the delivery capability to meet their customer needs. This paper will describe the interfaces between the 3 primary disciplines used in the design process; weights and sizing, trajectory, and structural analysis, as well as the approach each discipline employs to streamline their particular piece of the design process.

An overview on Bernese projects in planetary geodesy and deep-space orbit determination

NASA Astrophysics Data System (ADS)

Bertone, S.; Jaeggi, A.; Arnold, D.; Girardin, V.; Hosseini, A.; Desprats, W.; Inamdar, J.

2017-12-01

The Astronomical Institute of the University of Bern (AIUB) is still a rather new player in the field of planetary geodesy and orbit determination using deep-space radio-tracking data. Nevertheless, our latest developments in the in-house Bernese GNSS Software (BSW) and the experience gained with the processing of GRAIL data opened the way to many research and collaboration opportunities. In this presentation, we give an overview on our current projects and advances, as well as on our ongoing collaborations. We will present closed-loop simulations of BepiColombo Mercury Planetary Orbiter (MPO) Doppler and altimetry data, including realistic noise models. We use our newly established simulation environment in the BSW and calibration results of the BepiColombo Laser Altimeter (BELA) performed by the Space Research and Planetary Sciences division of the University of Bern. The ultimate goal of these activities is to test different realistic scenarios of the BELA in-orbit performance to improve the recovery of Mercury geodesy and geophysical parameters. We recently started to work on the combined re-processing of all historical missions to Venus to improve their orbits and hence Venus gravity field using new available data (e.g., new atmospheric models), processing tools and techniques and computational power. We shall present our latest advances in processing Magellan data and towards a rigorous solution for the Venus gravity field, e.g., avoiding a step-wise processing as used by Konopliv et al. (1999). The AIUB is currently involved in the Joint Europa Mission proposal. In this framework we present our results for a realistic orbit and gravity field recovery based on simulated Doppler radio-tracking data from the planned scenario of a three months low altitude polar orbit around Europa. We describe our efforts in adapting our simulation tools to the peculiar environment of the Jovian satellite system. Eventually we briefly present the highlights of our latest results in Moon geodesy, including our latest gravity field and tidal parameters solutions from GRAIL data. A separate presentation will be dedicated to detail our Moon-related activities within this session.

2001 Mars Odyssey Project report

NASA Technical Reports Server (NTRS)

Spencer, D. A.; Gibbs, R. G.; Mase, R. A.; Plaut, J. J.; Saunders, R. S.

2002-01-01

The Mars Odyssey orbiter was launched on April 7, 2001, and arrived at Mars on October 24, 2001. The orbiter carries scientific instruments that will determine surface elemental composition, mineralogy and morphology, and measure the Mars radiation environment from orbit. In addition, the orbiter will serve as a data relay for future surface missions. This paper will present an overview of the Odyssey project, including the key elements of the spacecraft design, mission design and navigation, mission operations, and the science approach. The project's risk management process will be described. Initial findings of the science team will be summarized.

Preliminary Orbit Determination System (PODS) for Tracking and Data Relay Satellite System (TDRSS)-tracked target Spacecraft using the homotopy continuation method

NASA Technical Reports Server (NTRS)

Kirschner, S. M.; Samii, M. V.; Broaddus, S. R.; Doll, C. E.

1988-01-01

The Preliminary Orbit Determination System (PODS) provides early orbit determination capability in the Trajectory Computation and Orbital Products System (TCOPS) for a Tracking and Data Relay Satellite System (TDRSS)-tracked spacecraft. PODS computes a set of orbit states from an a priori estimate and six tracking measurements, consisting of any combination of TDRSS range and Doppler tracking measurements. PODS uses the homotopy continuation method to solve a set of nonlinear equations, and it is particularly effective for the case when the a priori estimate is not well known. Since range and Doppler measurements produce multiple states in PODS, a screening technique selects the desired state. PODS is executed in the TCOPS environment and can directly access all operational data sets. At the completion of the preliminary orbit determination, the PODS-generated state, along with additional tracking measurements, can be directly input to the differential correction (DC) process to generate an improved state. To validate the computational and operational capabilities of PODS, tests were performed using simulated TDRSS tracking measurements for the Cosmic Background Explorer (COBE) satellite and using real TDRSS measurements for the Earth Radiation Budget Satellite (ERBS) and the Solar Mesosphere Explorer (SME) spacecraft. The effects of various measurement combinations, varying arc lengths, and levels of degradation of the a priori state vector on the PODS solutions were considered.

Effect of ephemeris errors on the accuracy of the computation of the tangent point altitude of a solar scanning ray as measured by the SAGE 1 and 2 instruments

NASA Technical Reports Server (NTRS)

Buglia, James J.

1989-01-01

An analysis was made of the error in the minimum altitude of a geometric ray from an orbiting spacecraft to the Sun. The sunrise and sunset errors are highly correlated and are opposite in sign. With the ephemeris generated for the SAGE 1 instrument data reduction, these errors can be as large as 200 to 350 meters (1 sigma) after 7 days of orbit propagation. The bulk of this error results from errors in the position of the orbiting spacecraft rather than errors in computing the position of the Sun. These errors, in turn, result from the discontinuities in the ephemeris tapes resulting from the orbital determination process. Data taken from the end of the definitive ephemeris tape are used to generate the predict data for the time interval covered by the next arc of the orbit determination process. The predicted data are then updated by using the tracking data. The growth of these errors is very nearly linear, with a slight nonlinearity caused by the beta angle. An approximate analytic method is given, which predicts the magnitude of the errors and their growth in time with reasonable fidelity.

Enhanced orbit determination filter sensitivity analysis: Error budget development

NASA Technical Reports Server (NTRS)

Estefan, J. A.; Burkhart, P. D.

1994-01-01

An error budget analysis is presented which quantifies the effects of different error sources in the orbit determination process when the enhanced orbit determination filter, recently developed, is used to reduce radio metric data. The enhanced filter strategy differs from more traditional filtering methods in that nearly all of the principal ground system calibration errors affecting the data are represented as filter parameters. Error budget computations were performed for a Mars Observer interplanetary cruise scenario for cases in which only X-band (8.4-GHz) Doppler data were used to determine the spacecraft's orbit, X-band ranging data were used exclusively, and a combined set in which the ranging data were used in addition to the Doppler data. In all three cases, the filter model was assumed to be a correct representation of the physical world. Random nongravitational accelerations were found to be the largest source of error contributing to the individual error budgets. Other significant contributors, depending on the data strategy used, were solar-radiation pressure coefficient uncertainty, random earth-orientation calibration errors, and Deep Space Network (DSN) station location uncertainty.

Precise GPS orbits for geodesy

NASA Technical Reports Server (NTRS)

Colombo, Oscar L.

1994-01-01

The Global Positioning System (GPS) has become, in recent years, the main space-based system for surveying and navigation in many military, commercial, cadastral, mapping, and scientific applications. Better receivers, interferometric techniques (DGPS), and advances in post-processing methods have made possible to position fixed or moving receivers with sub-decimeter accuracies in a global reference frame. Improved methods for obtaining the orbits of the GPS satellites have played a major role in these achievements; this paper gives a personal view of the main developments in GPS orbit determination.

Radar cross section measurements of a scale model of the space shuttle orbiter vehicle

NASA Technical Reports Server (NTRS)

Yates, W. T.

1978-01-01

A series of microwave measurements was conducted to determine the radar cross section of the Space Shuttle Orbiter vehicle at a frequency and at aspect angles applicable to re-entry radar acquisition and tracking. The measurements were performed in a microwave anechoic chamber using a 1/15th scale model and a frequency applicable to C-band tracking radars. The data were digitally recorded and processed to yield statistical descriptions useful for prediction of orbiter re-entry detection and tracking ranges.

Gravitational orientation of the orbital complex, Salyut-6--Soyuz

NASA Technical Reports Server (NTRS)

Grecho, G. M.; Sarychev, V. A.; Legostayev, V. P.; Sazonov, V. V.; Gansvind, I. N.

1983-01-01

A simple mathematical model is proposed for the Salyut-6-Soyuz orbital complex motion with respect to the center of mass under the one-axis gravity-gradient orientation regime. This model was used for processing the measurements of the orbital complex motion parameters when the above orientation region was implemented. Some actual satellite motions are simulated and the satellite's aerodynamic parameters are determined. Estimates are obtained for the accuracy of measurements as well as that of the mathematical model.

GPS-Based Navigation and Orbit Determination for the AMSAT Phase 3D Satellite

NASA Technical Reports Server (NTRS)

Davis, George; Carpenter, Russell; Moreau, Michael; Bauer, Frank H.; Long, Anne; Kelbel, David; Martin, Thomas

2002-01-01

This paper summarizes the results of processing GPS data from the AMSAT Phase 3D (AP3) satellite for real-time navigation and post-processed orbit determination experiments. AP3 was launched into a geostationary transfer orbit (GTO) on November 16, 2000 from Kourou, French Guiana, and then was maneuvered into its HEO over the next several months. It carries two Trimble TANS Vector GPS receivers for signal reception at apogee and at perigee. Its spin stabilization mode currently makes it favorable to track GPS satellites from the backside of the constellation while at perigee, and to track GPS satellites from below while at perigee. To date, the experiment has demonstrated that it is feasible to use GPS for navigation and orbit determination in HEO, which will be of great benefit to planned and proposed missions that will utilize such orbits for science observations. It has also shown that there are many important operational considerations to take into account. For example, GPS signals can be tracked above the constellation at altitudes as high as 58000 km, but sufficient amplification of those weak signals is needed. Moreover, GPS receivers can track up to 4 GPS satellites at perigee while moving as fast as 9.8 km/sec, but unless the receiver can maintain lock on the signals long enough, point solutions will be difficult to generate. The spin stabilization of AP3, for example, appears to cause signal levels to fluctuate as other antennas on the satellite block the signals. As a result, its TANS Vectors have been unable to lock on to the GPS signals long enough to down load the broadcast ephemeris and then generate position and velocity solutions. AP3 is currently in its eclipse season, and thus most of the spacecraft subsystems have been powered off. In Spring 2002, they will again be powered up and AP3 will be placed into a three-axis stabilization mode. This will significantly enhance the likelihood that point solutions can be generated, and perhaps more important, that the receiver clock can be synchronized to GPS time. This is extremely important for real-time and post-processed orbit determination, where removal of receiver clock bias from the data time tags is needed, for time-tagging of science observations. Current analysis suggests that the inability to generate point solutions has allowed the TANS Vector clock bias to drift freely, being perhaps as large as 5-7 seconds by October, 2001, thus causing up to 50 km of along-track orbit error. The data collected in May, 2002 while in three-axis stabilized mode should provide a significant improvement in the orbit determination results.

A Southern Hemisphere radar meteor orbit survey

NASA Technical Reports Server (NTRS)

Baggaley, W. Jack; Steel, Duncan I.; Taylor, Andrew D.

1992-01-01

A meteor radar system has been operated on a routine basis near Christchurch, New Zealand, to determine the orbits of Earth-impacting interplanetary dust and meteoroids. The system sensitivity is +13 visual magnitude, corresponding to approximately 100 micron sized meteoroids. With an orbital precision of 2 degrees in angular elements and 10 percent in orbital energy (1/a), the operation yields an average of 1500 orbits daily with a total to date in excess of 10(exp 5). The use of pc's and automated data reduction permit the large orbital data sets we collect to be routinely reduced. Some illustrative examples are presented of the signal formats/processing and the results of data reduction, giving the individual orbital elements and hence the overall distributions. Current studies include the distribution of dust in the inner solar system; the influx of meteoroids associated with near-Earth asteroids; and the orbital structure existing in comet-produced streams.

First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations

NASA Technical Reports Server (NTRS)

Kahle, R.; Kazeminejad, B.; Kirschner, M.; Yoon, Y.; Kiehling, R.; D'Amico, S.

2007-01-01

TerraSAR-X is an advanced synthetic aperture radar satellite system for scientific and commercial applications that is realized in a public-private partnership between the German Aerospace Center (DLR) and the Astrium GmbH. TerraSAR-X was launched at June 15, 2007 on top of a Russian DNEPR-1 rocket into a 514 km sun-synchronous dusk-dawn orbit with an 11-day repeat cycle and will be operated for a period of at least 5 years during which it will provide high resolution SAR-data in the X-band. Due to the objectives of the interferometric campaigns the satellite has to comply to tight orbit control requirements, which are formulated in the form of a 250 m toroidal tube around a pre-flight determined reference trajectory (see [1] for details). The acquisition of the reference orbit was one of the main and key activities during the Launch and Early Orbit Phase (LEOP) and had to compensate for both injection errors and spacecraft safe mode attitude control thruster activities. The paper summarizes the activities of GSOC flight dynamics team during both LEOP and early Commissioning Phase, where the main tasks have been 1) the first-acquisition support via angle-tracking and GPS-based orbit determination, 2) maneuver planning for target orbit acquisition and maintenance, and 3) precise orbit and attitude determination for SAR processing support. Furthermore, a presentation on the achieved results and encountered problems will be addressed.

Orbit determination and orbit control for the Earth Observing System (EOS) AM spacecraft

NASA Technical Reports Server (NTRS)

Herberg, Joseph R.; Folta, David C.

1993-01-01

Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.

Near-real time orbit determination for the GPS, CHAMP, GRACE, TerraSAR-X, and TanDEM-X satellites

NASA Astrophysics Data System (ADS)

Michalak, Grzegorz; Koenig, Rolf

The GFZ German Research Centre for Geosciences developed a near-real time (NRT) orbit gen-eration system for GPS and Low Earth Orbiting (LEO) satellites to support radio occultation data processing for the CHAMP, GRACE, Terra-SAR-X and the upcoming TanDEM-X mis-sions and fast baseline determination for the TanDEM-X mission. Precise NRT orbits are being generated for the CHAMP and GRACE-A satellites since August 2006 and for TerraSAR-X since August 2007. For each LEO, the system consists of three independent chains delivering NRT orbits with different latencies and accuracies. The first chain generates in a preceding step NRT GPS orbits and clock biases and based thereon LEO orbits with delays of 30 minutes counted from the last measurement point to the time the orbit product is available. The orbit accuracies can be assessed via Satellite Laser Ranging (SLR) to 7 cm. The second chain is based on predicted GPS orbits from the International GNSS Service (IGS) but endowed with in-house estimated clock biases. This chain generates orbits with the same latency of 30 minutes but with better accuracies of 5 cm SLR RMS. The third chain, the least accurate but the fastest, is based on predicted IGS GPS orbits and clocks and delivers LEO orbits with latencies of 13 minutes and accuracies of 10 cm SLR RMS. The system design is such that it can easily be extended to cope with new satellites like TanDEM-X requiring precise and fast available orbits.

On-board orbit determination for low thrust LEO-MEO transfer by Consider Kalman Filtering and multi-constellation GNSS

NASA Astrophysics Data System (ADS)

Menzione, Francesco; Renga, Alfredo; Grassi, Michele

2017-09-01

In the framework of the novel navigation scenario offered by the next generation satellite low thrust autonomous LEO-to-MEO orbit transfer, this study proposes and tests a GNSS based navigation system aimed at providing on-board precise and robust orbit determination strategy to override rising criticalities. The analysis introduces the challenging design issues to simultaneously deal with the variable orbit regime, the electric thrust control and the high orbit GNSS visibility conditions. The Consider Kalman Filtering approach is here proposed as the filtering scheme to process the GNSS raw data provided by a multi-antenna/multi-constellation receiver in presence of uncertain parameters affecting measurements, actuation and spacecraft physical properties. Filter robustness and achievable navigation accuracy are verified using a high fidelity simulation of the low-thrust rising scenario and performance are compared with the one of a standard Extended Kalman Filtering approach to highlight the advantages of the proposed solution. Performance assessment of the developed navigation solution is accomplished for different transfer phases.

Solar Radiation Pressure Estimation and Analysis of a GEO Class of High Area-to-Mass Ratio Debris Objects

NASA Technical Reports Server (NTRS)

Kelecy, Tom; Payne, Tim; Thurston, Robin; Stansbery, Gene

2007-01-01

A population of deep space objects is thought to be high area-to-mass ratio (AMR) debris having origins from sources in the geosynchronous orbit (GEO) belt. The typical AMR values have been observed to range anywhere from 1's to 10's of m(sup 2)/kg, and hence, higher than average solar radiation pressure effects result in long-term migration of eccentricity (0.1-0.6) and inclination over time. However, the nature of the debris orientation-dependent dynamics also results time-varying solar radiation forces about the average which complicate the short-term orbit determination processing. The orbit determination results are presented for several of these debris objects, and highlight their unique and varied dynamic attributes. Estimation or the solar pressure dynamics over time scales suitable for resolving the shorter term dynamics improves the orbit estimation, and hence, the orbit predictions needed to conduct follow-up observations.

Low-degree gravity change from GPS data of COSMIC and GRACE satellite missions

NASA Astrophysics Data System (ADS)

Lin, Tingjung; Hwang, Cheinway; Tseng, Tzu-Pang; Chao, B. F.

2012-01-01

This paper demonstrates estimation of time-varying gravity harmonic coefficients from GPS data of COSMIC and GRACE satellite missions. The kinematic orbits of COSMIC and GRACE are determined to the cm-level accuracy. The NASA Goddard's GEODYN II software is used to model the orbit dynamics of COSMIC and GRACE, including the effect of a static gravity field. The surface forces are estimated per one orbital period. Residual orbits generated from kinematic and reference orbits serve as observables to determine the harmonic coefficients in the weighted-constraint least-squares. The monthly COSMIC and GRACE GPS data from September 2006 to December 2007 (16 months) are processed to estimate harmonic coefficients to degree 5. The geoid variations from the GPS and CSR RL04 (GRACE) solutions show consistent patterns over space and time, especially in regions of active hydrological changes. The monthly GPS-derived second zonal coefficient closely resembles the SLR-derived and CSR RL04 values, and third and fourth zonal coefficients resemble the CSR RL04 values.

Lunar far side surface navigation using Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON)

NASA Astrophysics Data System (ADS)

Hesar, Siamak G.; Parker, Jeffrey S.; Leonard, Jason M.; McGranaghan, Ryan M.; Born, George H.

2015-12-01

We study the application of Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) to track vehicles on the far side of the lunar surface. The LiAISON architecture is demonstrated to achieve accurate orbit determination solutions for various mission scenarios in the Earth-Moon system. Given the proper description of the force field, LiAISON is capable of producing absolute orbit determination solutions using relative satellite-to-satellite tracking observations alone. The lack of direct communication between Earth-based tracking stations and the far side of the Moon provides an ideal opportunity for implementing LiAISON. This paper presents a novel approach to use the LiAISON architecture to perform autonomous navigation of assets on the lunar far side surface. Relative measurements between a spacecraft placed in an EML-2 halo orbit and lunar surface asset(s) are simulated and processed. Comprehensive simulation results show that absolute states of the surface assets are observable with an achieved accuracy of the position estimate on the order of tens of meters.

Many-core computing for space-based stereoscopic imaging

NASA Astrophysics Data System (ADS)

McCall, Paul; Torres, Gildo; LeGrand, Keith; Adjouadi, Malek; Liu, Chen; Darling, Jacob; Pernicka, Henry

The potential benefits of using parallel computing in real-time visual-based satellite proximity operations missions are investigated. Improvements in performance and relative navigation solutions over single thread systems can be achieved through multi- and many-core computing. Stochastic relative orbit determination methods benefit from the higher measurement frequencies, allowing them to more accurately determine the associated statistical properties of the relative orbital elements. More accurate orbit determination can lead to reduced fuel consumption and extended mission capabilities and duration. Inherent to the process of stereoscopic image processing is the difficulty of loading, managing, parsing, and evaluating large amounts of data efficiently, which may result in delays or highly time consuming processes for single (or few) processor systems or platforms. In this research we utilize the Single-Chip Cloud Computer (SCC), a fully programmable 48-core experimental processor, created by Intel Labs as a platform for many-core software research, provided with a high-speed on-chip network for sharing information along with advanced power management technologies and support for message-passing. The results from utilizing the SCC platform for the stereoscopic image processing application are presented in the form of Performance, Power, Energy, and Energy-Delay-Product (EDP) metrics. Also, a comparison between the SCC results and those obtained from executing the same application on a commercial PC are presented, showing the potential benefits of utilizing the SCC in particular, and any many-core platforms in general for real-time processing of visual-based satellite proximity operations missions.

Monte Carlo Analysis as a Trajectory Design Driver for the Transiting Exoplanet Survey Satellite (TESS) Mission

NASA Technical Reports Server (NTRS)

Nickel, Craig; Parker, Joel; Dichmann, Don; Lebois, Ryan; Lutz, Stephen

2016-01-01

The Transiting Exoplanet Survey Satellite (TESS) will be injected into a highly eccentric Earth orbit and fly 3.5 phasing loops followed by a lunar flyby to enter a mission orbit with lunar 2:1 resonance. Through the phasing loops and mission orbit, the trajectory is significantly affected by lunar and solar gravity. We have developed a trajectory design to achieve the mission orbit and meet mission constraints, including eclipse avoidance and a 30-year geostationary orbit avoidance requirement. A parallelized Monte Carlo simulation was performed to validate the trajectory after injecting common perturbations, including launch dispersions, orbit determination errors, and maneuver execution errors. The Monte Carlo analysis helped identify mission risks and is used in the trajectory selection process.

Precise Orbit Determination of BeiDou Navigation Satellite System

NASA Astrophysics Data System (ADS)

He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald

2013-04-01

China has been developing its own independent satellite navigation system since decades. Now the COMPASS system, also known as BeiDou, is emerging and gaining more and more interest and attention in the worldwide GNSS communities. The current regional BeiDou system is ready for its operational service around the end of 2012 with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit satellites (IGSO) and four Medium Earth orbit (MEO) satellites in operation. Besides the open service with positioning accuracy of around 10m which is free to civilian users, both precise relative positioning, and precise point positioning are demonstrated as well. In order to enhance the BeiDou precise positioning service, Precise Orbit Determination (POD) which is essential of any satellite navigation system has been investigated and studied thoroughly. To further improving the orbits of different types of satellites, we study the impact of network coverage on POD data products by comparing results from tracking networks over the Chinese territory, Asian-Pacific, Asian and of global scale. Furthermore, we concentrate on the improvement of involving MEOs on the orbit quality of GEOs and IGSOs. POD with and without MEOs are undertaken and results are analyzed. Finally, integer ambiguity resolution which brings highly improvement on orbits and positions with GPS data is also carried out and its effect on POD data products is assessed and discussed in detail. Seven weeks of BeiDou data from a ground tracking network, deployed by Wuhan University is employed in this study. The test constellation includes four GEO, five IGSO and two MEO satellites in operation. The three-day solution approach is employed to enhance its strength due to the limited coverage of the tracking network and the small movement of most of the satellites. A number of tracking scenarios and processing schemas are identified and processed and overlapping orbit differences are utilized to qualify the estimated orbits and clocks. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. For the current tracking network, deploying tracking stations on the eastern side, for example in New Zealand and/or in Hawaii, will significantly reduce along-track biases of GEOs on the same side. The involvement of MEOs and ambiguity-fixing also make the orbits better but rather moderate. Key words: BeiDou, precise orbit determination (POD), tracking network, ambiguity-fixing

Time determination for spacecraft users of the Navstar Global Positioning System /GPS/

NASA Technical Reports Server (NTRS)

Grenchik, T. J.; Fang, B. T.

1977-01-01

Global Positioning System (GPS) navigation is performed by time measurements. A description is presented of a two body model of spacecraft motion. Orbit determination is the process of inferring the position, velocity, and clock offset of the user from measurements made of the user motion in the Newtonian coordinate system. To illustrate the effect of clock errors and the accuracy with which the user spacecraft time and orbit may be determined, a low-earth-orbit spacecraft (Seasat) as tracked by six Phase I GPS space vehicles is considered. The obtained results indicate that in the absence of unmodeled dynamic parameter errors clock biases may be determined to the nanosecond level. There is, however, a high correlation between the clock bias and the uncertainty in the gravitational parameter GM, i.e., the product of the universal gravitational constant and the total mass of the earth. It is, therefore, not possible to determine clock bias to better than 25 nanosecond accuracy in the presence of a gravitational error of one part per million.

The Importance of Semi-Major Axis Knowledge in the Determination of Near-Circular Orbits

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Schiesser, Emil R.

1998-01-01

Modem orbit determination has mostly been accomplished using Cartesian coordinates. This usage has carried over in recent years to the use of GPS for satellite orbit determination. The unprecedented positioning accuracy of GPS has tended to focus attention more on the system's capability to locate the spacecraft's location at a particular epoch than on its accuracy in determination of the orbit, per se. As is well-known, the latter depends on a coordinated knowledge of position, velocity, and the correlation between their errors. Failure to determine a properly coordinated position/velocity state vector at a given epoch can lead to an epoch state that does not propagate well, and/or may not be usable for the execution of orbit adjustment maneuvers. For the quite common case of near-circular orbits, the degree to which position and velocity estimates are properly coordinated is largely captured by the error in semi-major axis (SMA) they jointly produce. Figure 1 depicts the relationships among radius error, speed error, and their correlation which exist for a typical low altitude Earth orbit. Two familiar consequences are the relationship Figure 1 shows are the following: (1) downrange position error grows at the per orbit rate of 3(pi) times the SMA error; (2) a velocity change imparted to the orbit will have an error of (pi) divided by the orbit period times the SMA error. A less familiar consequence occurs in the problem of initializing the covariance matrix for a sequential orbit determination filter. An initial covariance consistent with orbital dynamics should be used if the covariance is to propagate well. Properly accounting for the SMA error of the initial state in the construction of the initial covariance accomplishes half of this objective, by specifying the partition of the covariance corresponding to down-track position and radial velocity errors. The remainder of the in-plane covariance partition may be specified in terms of the flight path angle error of the initial state. Figure 2 illustrates the effect of properly and not properly initializing a covariance. This figure was produced by propagating the covariance shown on the plot, without process noise, in a circular low Earth orbit whose period is 5828.5 seconds. The upper subplot, in which the proper relationships among position, velocity, and their correlation has been used, shows overall error growth, in terms of the standard deviations of the inertial position coordinates, of about half of the lower subplot, whose initial covariance was based on other considerations.

Experimental Study on the Precise Orbit Determination of the BeiDou Navigation Satellite System

PubMed Central

He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald

2013-01-01

The regional service of the Chinese BeiDou satellite navigation system is now in operation with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Besides the standard positioning service with positioning accuracy of about 10 m, both precise relative positioning and precise point positioning are already demonstrated. As is well known, precise orbit and clock determination is essential in enhancing precise positioning services. To improve the satellite orbits of the BeiDou regional system, we concentrate on the impact of the tracking geometry and the involvement of MEOs, and on the effect of integer ambiguity resolution as well. About seven weeks of data collected at the BeiDou Experimental Test Service (BETS) network is employed in this experimental study. Several tracking scenarios are defined, various processing schemata are designed and carried out; and then, the estimates are compared and analyzed in detail. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. The involvement of MEOs and ambiguity-fixing also make the orbits better. PMID:23529116

Experimental study on the precise orbit determination of the BeiDou navigation satellite system.

PubMed

He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald

2013-03-01

The regional service of the Chinese BeiDou satellite navigation system is now in operation with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit (IGSO) satellites and four Medium Earth Orbit (MEO) satellites. Besides the standard positioning service with positioning accuracy of about 10 m, both precise relative positioning and precise point positioning are already demonstrated. As is well known, precise orbit and clock determination is essential in enhancing precise positioning services. To improve the satellite orbits of the BeiDou regional system, we concentrate on the impact of the tracking geometry and the involvement of MEOs, and on the effect of integer ambiguity resolution as well. About seven weeks of data collected at the BeiDou Experimental Test Service (BETS) network is employed in this experimental study. Several tracking scenarios are defined, various processing schemata are designed and carried out; and then, the estimates are compared and analyzed in detail. The results show that GEO orbits, especially the along-track component, can be significantly improved by extending the tracking network in China along longitude direction, whereas IGSOs gain more improvement if the tracking network extends in latitude. The involvement of MEOs and ambiguity-fixing also make the orbits better.

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.g. improvements of the box-wing models, have been done to deliver best possible orbit solutions for the satellite. Quality control of the CPOD orbits is done by validating them with independent orbit solutions provided by the Copernicus POD Quality Working Group. The cross-comparison of orbit solutions from different institutions is essential to monitor and to improve the orbit accuracy because for Sentinel-1 and -2 this is the only possibility to externally assess the quality of the orbits. Sentinel-3 orbits may additionally be validated by using SLR and DORIS observations. This paper presents the Copernicus POD Service in terms of operations and orbital accuracy achieved by the different orbit products of the different missions. For Sentinel-1 and Sentinel-2, this paper presents the impact of the box-wing models. For Sentinel-3, the orbital accuracy will be assessed using the very first data after launch.

The tidal disruption of a star by a massive black hole

NASA Technical Reports Server (NTRS)

Evans, Charles R.; Kochanek, Christopher S.

1989-01-01

Results are reported from a three-dimensional numerical calculation of the tidal disruption of a low-mass main-sequence star on a parabolic orbit around a massive black hole (Mh = 10 to the 6th stellar mass). The postdisruption evolution is followed until hydrodynamic forces becomes negligible and the liberated gas becomes ballistic. Also given is the rate at which bound mass returns to pericenter after orbiting the hole once. The processes that determine the time scale to circularize the debris orbits and allow an accretion torus to form are discussed. This time scale and the time scales for radiative cooling and accretion inflow determine the onset and duration of the subsequent flare in the AGN luminosity.

Orbit Selection for Earth Observation Missions

NASA Technical Reports Server (NTRS)

King, J. C.

1978-01-01

The orbit selection process is simplified for most earth-oriented satellite missions by a restriction to circular orbits, which reduces the primary orbit characteristics to be determined to only two: altitude and inclination. A number of important mission performance characteristics depend on these choices, however, so a major part of the orbit selection task is concerned with developing the correlating relationships in clear and convenient forms to provide a basis for rational orbit selection procedures. The present approach to that task is organized around two major areas of mission performance, orbit plane precession and coverage pattern development, whose dependence on altitude and inclination is delineated graphically in design chart form. These charts provide a visual grasp of the relationships between the quantities cited above, as well as other important mission performance parameters including viewing time of day (solar), sensor swath width (and fields of view), swath sequencing, and pattern repeat condition and repeat periods.

Effect of Numerical Error on Gravity Field Estimation for GRACE and Future Gravity Missions

NASA Astrophysics Data System (ADS)

McCullough, Christopher; Bettadpur, Srinivas

2015-04-01

In recent decades, gravity field determination from low Earth orbiting satellites, such as the Gravity Recovery and Climate Experiment (GRACE), has become increasingly more effective due to the incorporation of high accuracy measurement devices. Since instrumentation quality will only increase in the near future and the gravity field determination process is computationally and numerically intensive, numerical error from the use of double precision arithmetic will eventually become a prominent error source. While using double-extended or quadruple precision arithmetic will reduce these errors, the numerical limitations of current orbit determination algorithms and processes must be accurately identified and quantified in order to adequately inform the science data processing techniques of future gravity missions. The most obvious numerical limitation in the orbit determination process is evident in the comparison of measured observables with computed values, derived from mathematical models relating the satellites' numerically integrated state to the observable. Significant error in the computed trajectory will corrupt this comparison and induce error in the least squares solution of the gravitational field. In addition, errors in the numerically computed trajectory propagate into the evaluation of the mathematical measurement model's partial derivatives. These errors amalgamate in turn with numerical error from the computation of the state transition matrix, computed using the variational equations of motion, in the least squares mapping matrix. Finally, the solution of the linearized least squares system, computed using a QR factorization, is also susceptible to numerical error. Certain interesting combinations of each of these numerical errors are examined in the framework of GRACE gravity field determination to analyze and quantify their effects on gravity field recovery.

Spin-up studies of the Space Shuttle Orbiter main gear tire

NASA Technical Reports Server (NTRS)

Daugherty, Robert H.; Stubbs, Sandy M.

1988-01-01

One of the factors needed to describe the wear behavior of the Space Shuttle Orbiter main gear tires is their behavior during the spin-up process. An experimental investigation of tire spin-up processes was conducted at the NASA Langley Research Center's Aircraft Landing Dynamics Facility. During the investigation, the influence of various parameters such as forward speed and sink speed on tire spin-up forces were evaluated. A mathematical model was developed to estimate drag forces and spin-up times and is presented. The effect of prerotation was explored and is discussed. Also included is a means of determining the sink speed of the orbiter at touchdown based upon the appearance of the rubber deposits left on the runway during spinup.

Anticipated results from dust experiments on cometary missions

NASA Technical Reports Server (NTRS)

Kissel, J.; Fechtig, H.; Grun, E.

1981-01-01

The major scientific objectives of a mission are: to determine the chemical nature and physical structure of comet nuclei, and to characterize the changes that occur as a function of time orbital position; to characterize the chemical and physical nature of the atmospheres and ionospheres of comets as well as the processes that occur in them, and to characterize the development of the atmospheres and ionospheres as functions of time and orbital position; and to determine the nature of comet tails and processes by which they are formed, and to characterize the interaction of comets with the solar wind. Since dust is a major constituent of a comet, the achievement of these goals requires the intensive study of the paticulate emission from a comet.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Barnes, Jason W., E-mail: jwbarnes@uidaho.ed

Main-sequence stars earlier than spectral-type approxF6 or so are expected to rotate rapidly due to their radiative exteriors. This rapid rotation leads to an oblate stellar figure. It also induces the photosphere to be hotter (by up to several thousand kelvin) at the pole than at the equator as a result of a process called gravity darkening that was first predicted by von Zeipel. Transits of extrasolar planets across such a non-uniform, oblate disk yield unusual and distinctive lightcurves that can be used to determine the relative alignment of the stellar rotation pole and the planet orbit normal. This spin-orbitmore » alignment can be used to constrain models of planet formation and evolution. Orderly planet formation and migration within a disk that is coplanar with the stellar equator will result in spin-orbit alignment. More violent planet-planet scattering events should yield spin-orbit misaligned planets. Rossiter-McLaughlin measurements of transits of lower-mass stars show that some planets are spin-orbit aligned, and some are not. Since Rossiter-McLaughlin measurements are difficult around rapid rotators, lightcurve photometry may be the best way to determine the spin-orbit alignment of planets around massive stars. The Kepler mission will monitor approx10{sup 4} of these stars within its sample. The lightcurves of any detected planets will allow us to probe the planet formation process around high-mass stars for the first time.« less

Precise orbit determination of Multi-GNSS constellation including GPS GLONASS BDS and GALIEO

NASA Astrophysics Data System (ADS)

Dai, Xiaolei

2014-05-01

In addition to the existing American global positioning system (GPS) and the Russian global navigation satellite system (GLONASS), the new generation of GNSS is emerging and developing, such as the Chinese BeiDou satellite navigation system (BDS) and the European GALILEO system. Multi-constellation is expected to contribute to more accurate and reliable positioning and navigation service. However, the application of multi-constellation challenges the traditional precise orbit determination (POD) strategy that was designed usually for single constellation. In this contribution, we exploit a more rigorous multi-constellation POD strategy for the ongoing IGS multi-GNSS experiment (MGEX) where the common parameters are identical for each system, and the frequency- and system-specified parameters are employed to account for the inter-frequency and inter-system biases. Since the authorized BDS attitude model is not yet released, different BDS attitude model are implemented and their impact on orbit accuracy are studied. The proposed POD strategy was implemented in the PANDA (Position and Navigation Data Analyst) software and can process observations from GPS, GLONASS, BDS and GALILEO together. The strategy is evaluated with the multi-constellation observations from about 90 MGEX stations and BDS observations from the BeiDou experimental tracking network (BETN) of Wuhan University (WHU). Of all the MGEX stations, 28 stations record BDS observation, and about 80 stations record GALILEO observations. All these data were processed together in our software, resulting in the multi-constellation POD solutions. We assessed the orbit accuracy for GPS and GLONASS by comparing our solutions with the IGS final orbit, and for BDS and GALILEO by overlapping our daily orbit solution. The stability of inter-frequency bias of GLONASS and inter-system biases w.r.t. GPS for GLONASS, BDS and GALILEO were investigated. At last, we carried out precise point positioning (PPP) using the multi-constellation POD orbit and clock products, and analyzed the contribution of these POD products to PPP. Keywords: Multi-GNSS, Precise Orbit Determination, Inter-frequency bias, Inter-system bias, Precise Point Positioning

Payload/cargo processing at the launch site

NASA Technical Reports Server (NTRS)

Ragusa, J. M.

1983-01-01

Payload processing at Kennedy Space Center is described, with emphasis on payload contamination control. Support requirements are established after documentation of the payload. The processing facilities feature enclosed, environmentally controlled conditions, with account taken of the weather conditions, door openings, accessing the payload, industrial activities, and energy conservation. Apparatus are also available for purges after Orbiter landing. The payloads are divided into horizontal, vertical, mixed, and life sciences and Getaway Special categories, which determines the processing route through the facilities. A canister/transport system features sealed containers for moving payloads from one facility building to another. All payloads are exposed to complete Orbiter bay interface checkouts in a simulator before actually being mounted in the bay.

An Empirical Method for Determining the Lunar Gravity Field. Ph.D. Thesis - George Washington Univ.

NASA Technical Reports Server (NTRS)

Ferrari, A. J.

1971-01-01

A method has been devised to determine the spherical harmonic coefficients of the lunar gravity field. This method consists of a two-step data reduction and estimation process. In the first step, a weighted least-squares empirical orbit determination scheme is applied to Doppler tracking data from lunar orbits to estimate long-period Kepler elements and rates. Each of the Kepler elements is represented by an independent function of time. The long-period perturbing effects of the earth, sun, and solar radiation are explicitly modeled in this scheme. Kepler element variations estimated by this empirical processor are ascribed to the non-central lunar gravitation features. Doppler data are reduced in this manner for as many orbits as are available. In the second step, the Kepler element rates are used as input to a second least-squares processor that estimates lunar gravity coefficients using the long-period Lagrange perturbation equations.

Extracting More Information from Passive Optical Tracking Observations for Reliable Orbit Element Generation

NASA Astrophysics Data System (ADS)

Bennett, J.; Gehly, S.

2016-09-01

This paper presents results from a preliminary method for extracting more orbital information from low rate passive optical tracking data. An improvement in the accuracy of the observation data yields more accurate and reliable orbital elements. A comparison between the orbit propagations from the orbital element generated using the new data processing method is compared with the one generated from the raw observation data for several objects. Optical tracking data collected by EOS Space Systems, located on Mount Stromlo, Australia, is fitted to provide a new orbital element. The element accuracy is determined from a comparison between the predicted orbit and subsequent tracking data or reference orbit if available. The new method is shown to result in a better orbit prediction which has important implications in conjunction assessments and the Space Environment Research Centre space object catalogue. The focus is on obtaining reliable orbital solutions from sparse data. This work forms part of the collaborative effort of the Space Environment Management Cooperative Research Centre which is developing new technologies and strategies to preserve the space environment (www.serc.org.au).

Model-based segmentation in orbital volume measurement with cone beam computed tomography and evaluation against current concepts.

PubMed

Wagner, Maximilian E H; Gellrich, Nils-Claudius; Friese, Karl-Ingo; Becker, Matthias; Wolter, Franz-Erich; Lichtenstein, Juergen T; Stoetzer, Marcus; Rana, Majeed; Essig, Harald

2016-01-01

Objective determination of the orbital volume is important in the diagnostic process and in evaluating the efficacy of medical and/or surgical treatment of orbital diseases. Tools designed to measure orbital volume with computed tomography (CT) often cannot be used with cone beam CT (CBCT) because of inferior tissue representation, although CBCT has the benefit of greater availability and lower patient radiation exposure. Therefore, a model-based segmentation technique is presented as a new method for measuring orbital volume and compared to alternative techniques. Both eyes from thirty subjects with no known orbital pathology who had undergone CBCT as a part of routine care were evaluated (n = 60 eyes). Orbital volume was measured with manual, atlas-based, and model-based segmentation methods. Volume measurements, volume determination time, and usability were compared between the three methods. Differences in means were tested for statistical significance using two-tailed Student's t tests. Neither atlas-based (26.63 ± 3.15 mm(3)) nor model-based (26.87 ± 2.99 mm(3)) measurements were significantly different from manual volume measurements (26.65 ± 4.0 mm(3)). However, the time required to determine orbital volume was significantly longer for manual measurements (10.24 ± 1.21 min) than for atlas-based (6.96 ± 2.62 min, p < 0.001) or model-based (5.73 ± 1.12 min, p < 0.001) measurements. All three orbital volume measurement methods examined can accurately measure orbital volume, although atlas-based and model-based methods seem to be more user-friendly and less time-consuming. The new model-based technique achieves fully automated segmentation results, whereas all atlas-based segmentations at least required manipulations to the anterior closing. Additionally, model-based segmentation can provide reliable orbital volume measurements when CT image quality is poor.

Effect of vacuum processing on outgassing within an orbiting molecular shield

NASA Technical Reports Server (NTRS)

Outlaw, R. A.

1982-01-01

The limiting hydrogen number density in an orbiting molecular shield is highly dependent on the outgassing rates from the materials of construction for the shield, experimental apparatus, and other hardware contained within the shield. Ordinary degassing temperatures used for ultrahigh vacuum studies (less than 450 C) are not sufficient to process metals so that the contribution to the number density within the shield due to outgassing is less than the theoretically attainable level (approximately 200 per cu. cm). Pure aluminum and type 347 stainless steel were studied as candidate shield materials. Measurements of their hydrogen concentration and diffusion coefficients were made, and the effects of high temperature vacuum processing (greater than 600 C) on their resulting outgassing rates was determined. The densities in a molecular shield due to the outgassing from either metal were substantially less ( 0.003) than the density due to the ambient atomic hydrogen flux at an orbital altitude of 500 km.

On the feasibility of phase only PPP for kinematic LEO orbits

NASA Astrophysics Data System (ADS)

Wallat, Christoph; Schön, Steffen

2016-04-01

Low Earth Orbiters (LEO) are satellites in altitudes up to 1000 kilometers. From the sensor data collected on board the Earth's gravity field can be recovered. Over the last 15 years several satellite missions were brought into space and the orbit determination improved over the years. To process the sensor data, precise positioning and timing of the satellite is mandatory. There are two approaches for precise orbit determination (POD) of LEO satellites. Kinematic orbits are based on GNSS observations and star camera data measured on board of the LEO. With a Precise Point Positioning (PPP) known from the terrestrial case, using ionospheric-free linear combinations P3 and L3 three-dimensional coordinates of the LEO can be estimated for every observation epoch. To counteract the challenges in kinematic orbit determination our approach is based on a technique called GNSS receiver clock modeling (RCM). Here the frequency stability of an external oscillator is used to model the behavior of the GNSS receiver clock with piecewise linear polynomials instead of estimating epoch-wise the receiver clock time offset as an unknown parameter. When using RCM the observation geometry is stabilized and the orbit coordinates and the receiver clock error can be estimated with a better precision. The satellites of the Gravity Recovery And Climate Experiment (GRACE) mission are equipped with Ultra Stable quartz Oscillators (USO). The USO frequency stability is used to correct the GRACE GPS receiver clock. Therefore, receiver clock modeling is feasible for polynomials with a length up to 60 seconds, leading to improved mean PDOP values of 30 % and smaller formal mean standard deviations of the coordinates between 6 and 33 %. We developed a new approach for GRACE orbits using kinematic PPP with clock modeling and tested our approach with simulated and real GPS data. The idea to use only carrier phase observations in the final processing and no code measurements leads to a reduced number of observations and changes in parameter correlation in the adjustment. Canceling the code observations out of the normal equation system is possible due to a technique named parameter lumping, which will be explained in detail. The estimated coordinates of our phase only approach are comparable to the conventional PPP solution concerning standard deviations and RMS values. We will point out the advantages of our approach for the kinematic orbit determination of the GRACE satellites also for improvements in computing phase ambiguities.

Martian Meteorology: Determination of Large Scale Weather Patterns from Surface Measurements

NASA Technical Reports Server (NTRS)

Murphy, James R.; Haberle, Robert M.; Bridger, Alison F. C.

1998-01-01

We employed numerical modelling of the martian atmosphere, and our expertise in understanding martian atmospheric processes, to better understand the coupling between lower and upper atmosphere processes. One practical application of this work has been our involvement with the ongoing atmospheric aerobraking which the Mars Global Surveyor (MGS) spacecraft is currently undergoing at Mars. Dr. Murphy is currently a member of the Mars Global Surveyor (MGS) Aerobraking Atmospheric Advisory Group (AAG). He was asked to participate in this activity based upon his knowledge of martian atmospheric dynamical processes. Aerobraking is a process whereby a spacecraft, in an elliptical orbit, passes through the upper layers of the atmosphere (in this instance Mars). This passage through the atmosphere 'drags'upon the spacecraft, gradually reducing its orbital velocity. This has the effect, over time, of converting the elliptical orbit to a circular orbit, which is the desired mapping orbit for MGS. Carrying out aerobraking eliminates the need for carrying large amounts of fuel on the spacecraft to execute an engine burn to achieve the desired orbit. Eliminating the mass of the fuel reduces the cost of launch. Damage to one of MGS's solar panels shortly after launch has resulted in a less aggressive extended in time aerobraking phase which will not end until March, 1999. Phase I extended from Sept. 1997 through March 1998. During this time period, Dr. Murphy participated almost daily in the AAG meetings, and beginning in December 1997 lead the meeting several times per week. The leader of each of the daily AAG meetings took the results of that meeting (current state of the atmosphere, identification of any time trends or spatial patterns in upper atmosphere densities, etc.) forward to the Aerobraking Planning Group (APG) meeting, at which time the decision was made to not change MGS orbit, to lower the orbit to reach higher densities (greater 'drag'), or raise the orbit to avoid experiencing excessive, possibly damaging densities.

Products of the SNPP VIIRS SD Screen Transmittance and the SD BRDFs From Both Yaw Maneuver and Regular On-Orbit Data

NASA Technical Reports Server (NTRS)

Lei, Ning; Xiong, Xiaoxiong

2017-01-01

To ensure data quality, the Earth-observing Visible Infrared Imaging Radiometer Suite (VIIRS) on the Suomi National Polar-orbiting Partnership satellite regularly performs on-orbit radiometric calibration of its 22 spectral bands. The primary radiance source for the calibration of the VIIRS reflective solar bands (RSBs) is a sunlit onboard solar diffuser (SD).During the calibration process, sunlight goes through a perforated plate (the SD screen) and then strikes the SD. The sunlight, scattered off the SD of near-Lambertian property, is used for the calibration. Consequently, the spectral radiance of the scattered sunlight is proportional to the product of the SD screen transmittance and the SD bidirectional reflectance distribution function (BRDF) value at the observation direction. The BRDF value is decomposed to the product of its initial value at launch and a numerical degradation factor that quantifies the decrease from the initial value. The degradation factor is determined by an onboard SD stability monitor (SDSM). During the BRDF degradation factor determination process, the SDSM receives the SD scattered sunlight and the sunlight that goes through another perforated plate at almost the same time. The ratio of the signal strengths from the two observations is used to determine the BRDF degradation factor. Consequently, the RSB radiometric calibration requires the accurate knowledge of the product of the SD screen transmittance and the initial BRDF value as sensed by the RSB and the SDSM detectors. We use both yaw maneuver and a small portion of regular on-orbit data to determine the products.

Optical navigation during the Voyager Neptune encounter

NASA Technical Reports Server (NTRS)

Riedel, J. E.; Owen, W. M., Jr.; Stuve, J. A.; Synnott, S. P.; Vaughan, R. M.

1990-01-01

Optical navigation techniques were required to successfully complete the planetary exploration phase of the NASA deep-space Voyager mission. The last of Voyager's planetary encounters, with Neptune, posed unique problems from an optical navigation standpoint. In this paper we briefly review general aspects of the optical navigation process as practiced during the Voyager mission, and discuss in detail particular features of the Neptune encounter which affected optical navigation. New approaches to the centerfinding problem were developed for both stars and extended bodies, and these are described. Results of the optical navigation data analysis are presented, as well as a description of the optical orbit determination system and results of its use during encounter. Partially as a result of the optical navigation processing, results of scientific significance were obtained. These results include the discovery and orbit determination of several new satellites of Neptune and the determination of the size of Triton, Neptune's largest moon.

Automated Construction of Molecular Active Spaces from Atomic Valence Orbitals.

PubMed

Sayfutyarova, Elvira R; Sun, Qiming; Chan, Garnet Kin-Lic; Knizia, Gerald

2017-09-12

We introduce the atomic valence active space (AVAS), a simple and well-defined automated technique for constructing active orbital spaces for use in multiconfiguration and multireference (MR) electronic structure calculations. Concretely, the technique constructs active molecular orbitals capable of describing all relevant electronic configurations emerging from a targeted set of atomic valence orbitals (e.g., the metal d orbitals in a coordination complex). This is achieved via a linear transformation of the occupied and unoccupied orbital spaces from an easily obtainable single-reference wave function (such as from a Hartree-Fock or Kohn-Sham calculations) based on projectors to targeted atomic valence orbitals. We discuss the premises, theory, and implementation of the idea, and several of its variations are tested. To investigate the performance and accuracy, we calculate the excitation energies for various transition-metal complexes in typical application scenarios. Additionally, we follow the homolytic bond breaking process of a Fenton reaction along its reaction coordinate. While the described AVAS technique is not a universal solution to the active space problem, its premises are fulfilled in many application scenarios of transition-metal chemistry and bond dissociation processes. In these cases the technique makes MR calculations easier to execute, easier to reproduce by any user, and simplifies the determination of the appropriate size of the active space required for accurate results.

Preliminary GPS orbit determination results for the Extreme Ultraviolet Explorer

NASA Technical Reports Server (NTRS)

Gold, Kenn; Bertiger, Willy; Wu, Sien; Yunck, Tom

1993-01-01

A single-frequency Motorola Global Positioning System (GPS) receiver was launched with the Extreme Ultraviolet Explorer mission in June 1992. The receiver utilizes dual GPS antennas placed on opposite sides of the satellite to obtain full GPS coverage as it rotates during its primary scanning mission. A data set from this GPS experiment has been processed at the Jet Propulsion Laboratory with the GIPSY-OASIS 2 software package. The single-frequency, dual antenna approach and the low altitude (approximately 500 km) orbit of the satellite create special problems for the GPS orbit determination analysis. The low orbit implies that the dynamics of the spacecraft will be difficult to model, and that atmospheric drag will be an important error source. A reduced-dynamic solution technique was investigated in which ad hoc accelerations were estimated at each time step to absorb dynamic model error. In addition, a single-frequency ionospheric correction was investigated, and a cycle-slip detector was written. Orbit accuracy is currently better than 5 m. Further optimization should improve this to about 1 m.

Space Trajectory Error Analysis Program (STEAP) for halo orbit missions. Volume 1: Analytic and user's manual

NASA Technical Reports Server (NTRS)

Byrnes, D. V.; Carney, P. C.; Underwood, J. W.; Vogt, E. D.

1974-01-01

Development, test, conversion, and documentation of computer software for the mission analysis of missions to halo orbits about libration points in the earth-sun system is reported. The software consisting of two programs called NOMNAL and ERRAN is part of the Space Trajectories Error Analysis Programs (STEAP). The program NOMNAL targets a transfer trajectory from Earth on a given launch date to a specified halo orbit on a required arrival date. Either impulsive or finite thrust insertion maneuvers into halo orbit are permitted by the program. The transfer trajectory is consistent with a realistic launch profile input by the user. The second program ERRAN conducts error analyses of the targeted transfer trajectory. Measurements including range, doppler, star-planet angles, and apparent planet diameter are processed in a Kalman-Schmidt filter to determine the trajectory knowledge uncertainty. Execution errors at injection, midcourse correction and orbit insertion maneuvers are analyzed along with the navigation uncertainty to determine trajectory control uncertainties and fuel-sizing requirements. The program is also capable of generalized covariance analyses.

DynAstVO : a Europlanet database of NEA orbits

NASA Astrophysics Data System (ADS)

Desmars, J.; Thuillot, W.; Hestroffer, D.; David, P.; Le Sidaner, P.

2017-09-01

DynAstVO is a new orbital database developed within the Europlanet 2020 RI and the Virtual European Solar and Planetary Access (VESPA) frameworks. The database is dedicated to Near-Earth asteroids and provide parameters related to orbits: osculating elements, observational information, ephemeris through SPICE kernel, and in particular, orbit uncertainty and associated covariance matrix. DynAstVO is daily updated on a automatic process of orbit determination on the basis of the Minor Planet Electronic Circulars that reports new observations or the discover of a new asteroid. This database conforms to EPN-TAP environment and is accessible through VO protocols and on the VESPA portal web access (http://vespa.obspm.fr/). A comparison with other classical databases such as Astorb, MPCORB, NEODyS and JPL is also presented.

The COLA Collision Avoidance Method

NASA Astrophysics Data System (ADS)

Assmann, K.; Berger, J.; Grothkopp, S.

2009-03-01

In the following we present a collision avoidance method named COLA. The method has been designed to predict collisions for Earth orbiting spacecraft on any orbits, including orbit changes, with other space-born objects. The point in time of a collision and the collision probability are determined. To guarantee effective processing the COLA method uses a modular design and is composed of several components which are either developed within this work or deduced from existing algorithms: A filtering module, the close approach determination, the collision detection and the collision probability calculation. A software tool which implements the COLA method has been verified using various test cases built from sample missions. This software has been implemented in the C++ programming language and serves as a universal collision detection tool at LSE Space Engineering & Operations AG.

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.

Precise Tracking of the Magellan and Pioneer Venus Orbiters by Same-Beam Interferometry. Part 2: Orbit Determination Analysis

NASA Technical Reports Server (NTRS)

Folkner, W. M.; Border, J. S.; Nandi, S.; Zukor, K. S.

1993-01-01

A new radio metric positioning technique has demonstrated improved orbit determination accuracy for the Magellan and Pioneer Venus Orbiter orbiters. The new technique, known as Same-Beam Interferometry (SBI), is applicable to the positioning of multiple planetary rovers, landers, and orbiters which may simultaneously be observed in the same beamwidth of Earth-based radio antennas. Measurements of carrier phase are differenced between spacecraft and between receiving stations to determine the plane-of-sky components of the separation vector(s) between the spacecraft. The SBI measurements complement the information contained in line-of-sight Doppler measurements, leading to improved orbit determination accuracy. Orbit determination solutions have been obtained for a number of 48-hour data arcs using combinations of Doppler, differenced-Doppler, and SBI data acquired in the spring of 1991. Orbit determination accuracy is assessed by comparing orbit solutions from adjacent data arcs. The orbit solution differences are shown to agree with expected orbit determination uncertainties. The results from this demonstration show that the orbit determination accuracy for Magellan obtained by using Doppler plus SBI data is better than the accuracy achieved using Doppler plus differenced-Doppler by a factor of four and better than the accuracy achieved using only Doppler by a factor of eighteen. The orbit determination accuracy for Pioneer Venus Orbiter using Doppler plus SBI data is better than the accuracy using only Doppler data by 30 percent.

Numerical Algorithms for Precise and Efficient Orbit Propagation and Positioning

NASA Astrophysics Data System (ADS)

Bradley, Ben K.

Motivated by the growing space catalog and the demands for precise orbit determination with shorter latency for science and reconnaissance missions, this research improves the computational performance of orbit propagation through more efficient and precise numerical integration and frame transformation implementations. Propagation of satellite orbits is required for astrodynamics applications including mission design, orbit determination in support of operations and payload data analysis, and conjunction assessment. Each of these applications has somewhat different requirements in terms of accuracy, precision, latency, and computational load. This dissertation develops procedures to achieve various levels of accuracy while minimizing computational cost for diverse orbit determination applications. This is done by addressing two aspects of orbit determination: (1) numerical integration used for orbit propagation and (2) precise frame transformations necessary for force model evaluation and station coordinate rotations. This dissertation describes a recently developed method for numerical integration, dubbed Bandlimited Collocation Implicit Runge-Kutta (BLC-IRK), and compare its efficiency in propagating orbits to existing techniques commonly used in astrodynamics. The BLC-IRK scheme uses generalized Gaussian quadratures for bandlimited functions. It requires significantly fewer force function evaluations than explicit Runge-Kutta schemes and approaches the efficiency of the 8th-order Gauss-Jackson multistep method. Converting between the Geocentric Celestial Reference System (GCRS) and International Terrestrial Reference System (ITRS) is necessary for many applications in astrodynamics, such as orbit propagation, orbit determination, and analyzing geoscience data from satellite missions. This dissertation provides simplifications to the Celestial Intermediate Origin (CIO) transformation scheme and Earth orientation parameter (EOP) storage for use in positioning and orbit propagation, yielding savings in computation time and memory. Orbit propagation and position transformation simulations are analyzed to generate a complete set of recommendations for performing the ITRS/GCRS transformation for a wide range of needs, encompassing real-time on-board satellite operations and precise post-processing applications. In addition, a complete derivation of the ITRS/GCRS frame transformation time-derivative is detailed for use in velocity transformations between the GCRS and ITRS and is applied to orbit propagation in the rotating ITRS. EOP interpolation methods and ocean tide corrections are shown to impact the ITRS/GCRS transformation accuracy at the level of 5 cm and 20 cm on the surface of the Earth and at the Global Positioning System (GPS) altitude, respectively. The precession-nutation and EOP simplifications yield maximum propagation errors of approximately 2 cm and 1 m after 15 minutes and 6 hours in low-Earth orbit (LEO), respectively, while reducing computation time and memory usage. Finally, for orbit propagation in the ITRS, a simplified scheme is demonstrated that yields propagation errors under 5 cm after 15 minutes in LEO. This approach is beneficial for orbit determination based on GPS measurements. We conclude with a summary of recommendations on EOP usage and bias-precession-nutation implementations for achieving a wide range of transformation and propagation accuracies at several altitudes. This comprehensive set of recommendations allows satellite operators, astrodynamicists, and scientists to make informed decisions when choosing the best implementation for their application, balancing accuracy and computational complexity.

The orbital evolution of the AMOR asteroidal group during 11,550 years

NASA Astrophysics Data System (ADS)

Babadzhanov, P. B.; Zausaev, A. F.; Pushkaryov, A. N.

The orbital evolution of twenty seven Amor asteroids was determined by the Everhart method for the time interval from 2250 AD to 9300 BC. Closest encounters with terrestrial planets are calculated in the evolution process. Stable resonances with Venus, Earth and Jupiter over the period from 2250 AD to 9300 BC have been obtained. Theoretical coordinates of radiants on initial and final moments of integrating were calculated.

Maintenance of the catalog of artificial objects in space.

NASA Astrophysics Data System (ADS)

Khutorovskij, Z. N.

1994-01-01

The catalog of artificial objects in space (AOS) is useful for estimating the safety of space flights, for constructing temporal and spatial models of the flux of AOS, for determining when and where dangerous AOS will break up, for tracking inoperative instruments and space stations, for eliminating false alarms that are triggered by observations of AOS in the Ballistic Missile Early Warning System and in the Anti-Missile system, etc. At present, the Space Surveillance System (located in the former USSR) automatically maintains a catalog consisting of more than 5000 AOS with dimensions of at least 10 cm. The orbital parameters are continuously updated from radar tracking data. The author describes the software which is used to process the information. He presents some of the features of the system itself, including the number of objects in various stages of the tracking process, the orbital parameters of AOS which break up, and how the fragments are detected, the accuracy of tracking and predicting the orbits of the AOS, and the accuracy with which we can estimate when and where an AOS will break up. As an example, the author presents the results of determination of the time when the orbiting complex Salyut-7 - Kosmos-1686 will break up, and where it will impact.

The Laser Ranging Experiment of the Lunar Reconnaissance Orbiter: Five Years of Operations and Data Analysis

NASA Technical Reports Server (NTRS)

Mao, Dandan; McGarry, Jan F.; Mazarico, Erwan; Neumann, Gregory A.; Sun, Xiaoli; Torrence, Mark H.; Zagwodzki, Thomas W.; Rowlands, David D.; Hoffman, Evan D.; Horvath, Julie E.;

Accuracy of estimating the masses of Phobos and Deimos from multiple Viking orbiter encounters

NASA Technical Reports Server (NTRS)

Tolson, R. H.; Mason, M. L.

1975-01-01

The problem was investigated of estimating the masses of Phobos and Deimos from Doppler and onboard optical measurements during the Viking extended mission. A Kalman filter was used to analyze the effects of gravitational uncertainties and nongravitational accelerations. These accelerations destroy the dynamical integrity of the orbit, and multibatch or limited memory filtering is preferred to single batch processing. Optical tracking is essential to improve the relative orbit geometry. The masses can be determined to about 10% and 25% respectively for Phobos and Deimos, assuming satellite densities of about 3 gr/cu cm.

A Stable Clock Error Model Using Coupled First and Second Order Gauss-Markov Processes

NASA Technical Reports Server (NTRS)

Carpenter, Russell; Lee, Taesul

2008-01-01

Long data outages may occur in applications of global navigation satellite system technology to orbit determination for missions that spend significant fractions of their orbits above the navigation satellite constellation(s). Current clock error models based on the random walk idealization may not be suitable in these circumstances, since the covariance of the clock errors may become large enough to overflow flight computer arithmetic. A model that is stable, but which approximates the existing models over short time horizons is desirable. A coupled first- and second-order Gauss-Markov process is such a model.

Analysis of HY2A precise orbit determination using DORIS

NASA Astrophysics Data System (ADS)

Gao, Fan; Peng, Bibo; Zhang, Yu; Evariste, Ngatchou Heutchi; Liu, Jihua; Wang, Xiaohui; Zhong, Min; Lin, Mingsen; Wang, Nazi; Chen, Runjing; Xu, Houze

2015-03-01

HY2A is the first Chinese marine dynamic environment satellite. The payloads include a radar altimeter to measure the sea surface height in combination with a high precision orbit to be determined from tracking data. Onboard satellite tracking includes GPS, SLR, and the DORIS DGXX receiver which delivers phase and pseudo-range measurements. CNES releases raw phase and pseudo-range measurements with RINEX DORIS 3.0 format and pre-processed Doppler range-rate with DORIS 2.2 data format. However, the VMSI software package developed by Van Martin Systems, Inc which is used to estimate HY2A DORIS orbits can only process Doppler range-rate but not the DORIS phase data which are available with much shorter latency. We have proposed a method of constructing the phase increment data, which are similar to range-rate data, from RINEX DORIS 3.0 phase data. We compute the HY2A orbits from June, 2013 to August, 2013 using the POD strategy described in this paper based on DORIS 2.2 range-rate data and our reconstructed phase increment data. The estimated orbits are evaluated by comparing with the CNES precise orbits and SLR residuals. Our DORIS-only orbits agree with the precise GPS + SLR + DORIS CNES orbits radially at 1-cm and about 3-cm in the other two directions. SLR test with the 50° cutoff elevation shows that the CNES orbit can achieve about 1.1-cm accuracy in radial direction and our DORIS-only POD solutions are slightly worse. In addition, other HY2A DORIS POD concerns are discussed in this paper. Firstly, we discuss the frequency offset values provided with the RINEX data and find that orbit accuracy for the case when the frequency offset is applied is worse than when it is not applied. Secondly, HY2A DORIS antenna z-offsets are estimated using two kinds of measurements from June, 2013 to August, 2013. The results show that the measurement errors contribute a total of about 2-cm difference of estimated z-offset. Finally, we estimate HY2A orbits selecting 3 days with severe geomagnetic storm activity and SLR residuals suggest that estimating a drag coefficient every 6 h without any constraint is sufficient for maintaining orbit accuracy.

Accuracy assessment of BDS precision orbit determination and the influence analysis of site distribution

NASA Astrophysics Data System (ADS)

Chen, Ming; Guo, Jiming; Li, Zhicai; Zhang, Peng; Wu, Junli; Song, Weiwei

2017-04-01

BDS precision orbit determination is a key content of the BDS application, but the inadequate ground stations and the poor distribution of the network are the main reasons for the low accuracy of BDS precise orbit determination. In this paper, the BDS precise orbit determination results are obtained by using the IGS MGEX stations and the Chinese national reference stations，the accuracy of orbit determination of GEO, IGSO and MEO is 10.3cm, 2.8cm and 3.2cm, and the radial accuracy is 1.6cm,1.9cm and 1.5cm.The influence of ground reference stations distribution on BDS precise orbit determination is studied. The results show that the Chinese national reference stations contribute significantly to the BDS orbit determination, the overlap precision of GEO/IGSO/MEO satellites were improved by 15.5%, 57.5% and 5.3% respectively after adding the Chinese stations.Finally, the results of ODOP(orbit distribution of precision) and SLR are verified. Key words: BDS precise orbit determination; accuracy assessment；Chinese national reference stations；reference stations distribution；orbit distribution of precision

Prediction, Measurement, and Control of Spacecraft Charging Hazards on the International Space Station(ISS)

NASA Astrophysics Data System (ADS)

Koontz, Steve; Alred, John; Ellison, Amy; Patton, Thomas; Minow, Joseph; Spetch, William

2010-09-01

Orbital inclination, 51.6 degrees, and altitude range, 300 to 400 km,(low-Earth orbit or LEO) determine the ISS spacecraft charging environment. Specific interactions of the ISS electrical power system and metallic structure with the Earth’s ionospheric plasma and the geomagnetic field dominate spacecraft charging processes for ISS. ISS also flies through auroral electron streams at high latitudes. In this paper, we report the character of ISS spacecraft charging processes in Earth’s ionosphere, the results of measurement and modelling of the subject charging processes, and the safety issues for ISS itself as well as for ISS interoperability with respect to extra vehicular activity(EVA) and visiting vehicle proximity operations.

Orbit-determination performance of Doppler data for interplanetary cruise trajectories. Part 2: 8.4-GHz performance and data-weighting strategies

NASA Technical Reports Server (NTRS)

Ulvestad, J. S.

1992-01-01

A consider error covariance analysis was performed in order to investigate the orbit-determination performance attainable using two-way (coherent) 8.4-GHz (X-band) Doppler data for two segments of the planned Mars Observer trajectory. The analysis includes the effects of the current level of calibration errors in tropospheric delay, ionospheric delay, and station locations, with particular emphasis placed on assessing the performance of several candidate elevation-dependent data-weighting functions. One weighting function was found that yields good performance for a variety of tracking geometries. This weighting function is simple and robust; it reduces the danger of error that might exist if an analyst had to select one of several different weighting functions that are highly sensitive to the exact choice of parameters and to the tracking geometry. Orbit-determination accuracy improvements that may be obtained through the use of calibration data derived from Global Positioning System (GPS) satellites also were investigated, and can be as much as a factor of three in some components of the spacecraft state vector. Assuming that both station-location errors and troposphere calibration errors are reduced simultaneously, the recommended data-weighting function need not be changed when GPS calibrations are incorporated in the orbit-determination process.

Optical Observations of Space Debris

NASA Technical Reports Server (NTRS)

Seitzer, Patrick; Abercromby, Kira; Rodriquez, Heather; Barker, Edwin S.; Kelecy, Thomas

2008-01-01

This viewgraph presentation reviews the use of optical telescopes to observe space debris. .It will present a brief review of how the survey is conducted, and what some of the significant results encompass. The goal is to characterize the population of debris objects at GEO, with emphasis on the faint object population. Because the survey observations extend over a very short arc (5 minutes), a full six parameter orbit can not be determined. Recently we have begun to use a second telescope, the 0.9-m at CTIO, as a chase telescope to do follow-up observations of potential GEO debris candidates found by MODEST. With a long enough sequence of observations, a full six-parameter orbit including eccentricity can be determined. The project has used STK since inception for planning observing sessions based on the distribution of bright cataloged objects and the anti-solar point (to avoid eclipse). Recently, AGI's Orbit Determination Tool Kit (ODTK) has been used to determine orbits, including the effects of solar radiation pressure. Since an unknown fraction of the faint debris at GEO has a high area-to-mass ratio (A/M), the orbits are perturbed significantly by solar radiation. The ODTK analysis results indicate that temporal variations in the solar perturbations, possibly due to debris orientation dynamics, can be estimated in the OD process. Additionally, the best results appear to be achieved when solar forces orthogonal to the object-Sun line are considered. Determining the A/M of individual objects and the distribution of A/M values of a large sample of debris is important to understanding the total population of debris at GEO

Optical Observations of GEO Debris with Two Telescopes

NASA Technical Reports Server (NTRS)

Seitzer, P.; Abercromby, K.; Rodriguez, H.; Barker, E.

2007-01-01

For several years, the Michigan Orbital DEbris Survey Telescope (MODEST), the University of Michigan s 0.6/0.9-m Schmidt telescope on Cerro Tololo Inter-American Observatory in Chile has been used to survey the debris population at GEO in the visible regime. Magnitudes, positions, and angular rates are determined for GEO objects as they move across the telescope s field-of-view (FOV) during a 5-minute window. This short window of time is not long enough to determine a full six parameter orbit so usually a circular orbit is assumed. A longer arc of time is necessary to determine eccentricity and to look for changes in the orbit with time. MODEST can follow objects in real-time, but only at the price of stopping survey operations. A second telescope would allow for longer arcs of orbit to obtain the full six orbital parameters, as well as assess the changes over time. An additional benefit of having a second telescope is the capability of obtaining BVRI colors of the faint targets, aiding efforts to determine the material type of faint debris. For 14 nights in March 2007, two telescopes were used simultaneously to observe the GEO debris field. MODEST was used exclusively in survey mode. As objects were detected, they were handed off in near real-time to the Cerro Tololo 0.9-m telescope for follow-up observations. The goal was to determine orbits and colors for all objects fainter than R = 15th magnitude (corresponds to 1 meter in size assuming a 0.2 albedo) detected by MODEST. The hand-off process was completely functional during the final eight nights and follow-ups for objects from night-to-night were possible. The cutoff magnitude level of 15th was selected on the basis of an abrupt change in the observed angular rate distribution in the MODEST surveys. Objects brighter than 15th magnitude tend to lie on a well defined locus in the angular rate plane (and have orbits in the catalog), while fainter objects fill the plane almost uniformly. We need to determine full six-parameter orbits to investigate what causes this change in observed angular rates. Are these faint objects either the same population of high area-to-mass (A/M) objects on eccentric orbits as discovered by the ESA Space Debris Telescope (Schildknecht, et al. 2004), or are they just normal debris from breakups in GEO?

Utilizing the Lunar Laser Ranging datasets alongside the radioscience data from the Lunar Reconnaissance Orbiter to improve the dynamical model of the Moon

NASA Astrophysics Data System (ADS)

Viswanathan, Vishnu; Fienga, Agnes; Laskar, Jacques; Manche, Herve; Torre, Jean-Marie; Courde, Clément; Exertier, Pierre

2015-08-01

In this poster we elaborate the use of raw navigation data (range and Doppler observations) from the Lunar Reconnaissance Orbiter (LRO) available on the Planetary Data System (PDS), in order to study the orbit of this probe using the orbit determination software (GINS) developed by the French space agency (CNES). The constraints that are derived from this process on combining with the high precision Lunar Laser Ranging (LLR) datasets which are spread over 40 years, facilitates an improved dynamical modeling of the Moon. In addition, the possible advantages that could be exploited by the LLR experiments when operated with lasers in the IR wavelength are analyzed.

Independent Orbiter Assessment (IOA): Analysis of the crew equipment subsystem

NASA Technical Reports Server (NTRS)

Sinclair, Susan; Graham, L.; Richard, Bill; Saxon, H.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical (PCIs) items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results coresponding to the Orbiter crew equipment hardware are documented. The IOA analysis process utilized available crew equipment hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 352 failure modes analyzed, 78 were determined to be PCIs.

Navigation for the new millennium: Autonomous navigation for Deep Space 1

NASA Technical Reports Server (NTRS)

Reidel, J. E.; Bhaskaran, S.; Synnott, S. P.; Desai, S. D.; Bollman, W. E.; Dumont, P. J.; Halsell, C. A.; Han, D.; Kennedy, B. M.; Null, G. W.;

Statistical Orbit Determination using the Particle Filter for Incorporating Non-Gaussian Uncertainties

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.

Atmospheric density determination using high-accuracy satellite GPS data

NASA Astrophysics Data System (ADS)

Tingling, R.; Miao, J.; Liu, S.

2017-12-01

Atmospheric drag is the main error source in the orbit determination and prediction of low Earth orbit (LEO) satellites, however, empirical models which are used to account for atmosphere often exhibit density errors around 15 30%. Atmospheric density determination thus become an important topic for atmospheric researchers. Based on the relation between atmospheric drag force and the decay of orbit semi-major axis, we derived atmospheric density along the trajectory of CHAMP with its Rapid Science Orbit (RSO) data. Three primary parameters are calculated, including the ratio of cross sectional area to mass, drag coefficient, and the decay of semi-major axis caused by atmospheric drag. We also analyzed the source of error and made a comparison between GPS-derived and reference density. Result on 2 Dec 2008 shows that the mean error of GPS-derived density can decrease from 29.21% to 9.20% when time span adopted on the process of computation increase from 10min to 50min. Result for the whole December indicates that when the time span meet the condition that the amplitude of the decay of semi-major axis is much greater than its standard deviation, then density precision of 10% can be achieved.

Independent Orbiter Assessment (IOA): Analysis of the orbital maneuvering system

NASA Technical Reports Server (NTRS)

Prust, C. D.; Paul, D. J.; Burkemper, V. J.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Orbital Maneuvering System (OMS) hardware are documented. The OMS provides the thrust to perform orbit insertion, orbit circularization, orbit transfer, rendezvous, and deorbit. The OMS is housed in two independent pods located one on each side of the tail and consists of the following subsystems: Helium Pressurization; Propellant Storage and Distribution; Orbital Maneuvering Engine; and Electrical Power Distribution and Control. The IOA analysis process utilized available OMS hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluted and analyzed for possible failure modes and effects. Criticality was asigned based upon the severity of the effect for each failure mode.

System Engineering Issues for Avionics Survival in the Space Environment

NASA Technical Reports Server (NTRS)

Pavelitz, Steven

1999-01-01

This paper examines how the system engineering process influences the design of a spacecraft's avionics by considering the space environment. Avionics are susceptible to the thermal, radiation, plasma, and meteoroids/orbital debris environments. The environment definitions for various spacecraft mission orbits (LEO/low inclination, LEO/Polar, MEO, HEO, GTO, GEO and High ApogeeElliptical) are discussed. NASA models and commercial software used for environment analysis are reviewed. Applicability of technical references, such as NASA TM-4527 "Natural Orbital Environment Guidelines for Use in Aerospace Vehicle Development" is discussed. System engineering references, such as the MSFC System Engineering Handbook, are reviewed to determine how the environments are accounted for in the system engineering process. Tools and databases to assist the system engineer and avionics designer in addressing space environment effects on avionics are described and usefulness assessed.

Near-parabolic comets observed in 2006-2010. The individualized approach to 1/a-determination and the new distribution of original and future orbits

NASA Astrophysics Data System (ADS)

Królikowska, Małgorzata; Dybczyński, Piotr A.

2013-10-01

Dynamics of a complete sample of small perihelion distance near-parabolic comets discovered in the years 2006-2010 are studied (i.e. of 22 comets of qosc < 3.1 au). First, osculating orbits are obtained after a very careful positional data inspection and processing, including where appropriate, the method of data partitioning for determination of pre- and post-perihelion orbit for tracking then its dynamical evolution. The non-gravitational acceleration in the motion is detected for 50 per cent of investigated comets, in a few cases for the first time. Different sets of non-gravitational parameters are determined from pre- and post-perihelion data for some of them. The influence of the positional data structure on the possibility of the detection of non-gravitational effects and the overall precision of orbit determination is widely discussed. Secondly, both original and future orbits were derived by means of numerical integration of swarms of virtual comets obtained using a Monte Carlo cloning method. This method allows us to follow the uncertainties of orbital elements at each step of dynamical evolution. The complete statistics of original and future orbits that includes significantly different uncertainties of 1/a-values is presented, also in the light of our results obtained earlier. Basing on 108 comets examined by us so far, we conclude that only one of them, C/2007 W1 Boattini, seems to be a serious candidate for an interstellar comet. We also found that 53 per cent of 108 near-parabolic comets escaping in the future from the Solar system, and the number of comets leaving the Solar system as so called Oort spike comets (i.e. comets suffering very small planetary perturbations) is 14 per cent. A new method for cometary orbit quality assessment is also proposed by means of modifying the original method, introduced by Marsden, Sekanina & Everhart. This new method leads to a better diversification of orbit quality classes for contemporary comets.

Precise satellite orbit determination with particular application to ERS-1

NASA Astrophysics Data System (ADS)

Fernandes, Maria Joana Afonso Pereira

The motivation behind this study is twofold. First to assess the accuracy of ERS-1 long arc ephemerides using state of the art models. Second, to develop improved methods for determining precise ERS-1 orbits using either short or long arc techniques. The SATAN programs, for the computation of satellite orbits using laser data were used. Several facilities were added to the original programs: the processing of PRARE range and altimeter data, and a number of algorithms that allow more flexible solutions by adjusting a number of additional parameters. The first part of this study, before the launch of ERS-1, was done with SEAS AT data. The accuracy of SEASAT orbits computed with PRARE simulated data has been determined. The effect of temporal distribution of tracking data along the arc and the extent to which altimetry can replace range data have been investigated. The second part starts with the computation of ERS-1 long arc solutions using laser data. Some aspects of modelling the two main forces affecting ERS-l's orbit are investigated. With regard to the gravitational forces, the adjustment of a set of geopotential coefficients has been considered. With respect to atmospheric drag, extensive research has been carried out on determining the influence on orbit accuracy of the measurements of solar fluxes (P10.7 indices) and geomagnetic activity (Kp indices) used by the atmospheric model in the computation of atmospheric density at satellite height. Two new short arc methods have been developed: the Constrained and the Bayesian method. Both methods are dynamic and consist of solving for the 6 osculating elements. Using different techniques, both methods overcome the problem of normal matrix ill- conditioning by constraining the solution. The accuracy and applicability of these methods are discussed and compared with the traditional non-dynamic TAR method.

Orbit determination of the Sentinel satellites - preparations for GPS L2C-tracking

NASA Astrophysics Data System (ADS)

Peter, Heike; Fernández, Jaime; Fernández, Carlos; Féménias, Pierre

2017-04-01

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. Since April 2014 four Sentinel satellites have been launched (1A, 2A, 3A, and 1B). Sentinel-2B is expected to be launched in March 2017. Thus the CPOD Service will be operating five satellites simultaneously in spring 2017. 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 are additionally equipped with a laser retro reflector for Satellite Laser Ranging and a receiver for DORIS tracking. This allows an additional external validation of the Sentinel-3 orbit accuracy. 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. The main quality control of the CPOD orbits is done by validating them with independent orbit solutions provided by the Copernicus POD Quality Working Group. The cross-comparison of orbit solutions from different institutions is essential to monitor and to improve the orbit accuracy. The GPS receivers on the B-satellites have the capability to track L2C signal. The option is, however, not yet activated, because if enabled the old L2 signal can no longer be tracked by the receiver. The measurements of many old GPS IIA and IIR satellites would have to be discarded because of the missing second frequency. To be prepared for the future, tests and simulations are foreseen to learn about the impact of the new observable on the POD results. This paper presents the Copernicus POD Service in terms of operations and orbital accuracy achieved by the different orbit products of the different missions. The long-term evolution and progress of the service is presented and the impact and challenges following a future switch to L2C tracking are analysed.

Ground Optical Signal Processing Architecture for Contributing SSA Space Based Sensor Data

NASA Astrophysics Data System (ADS)

Koblick, D.; Klug, M.; Goldsmith, A.; Flewelling, B.; Jah, M.; Shanks, J.; Piña, R.

2014-09-01

The main objective of the DARPA program Orbit Outlook (O^2) is to improve the metric tracking and detection performance of the Space Situational Network (SSN) by adding a diverse low-cost network of contributing sensors to the Space Situational Awareness (SSA) mission. In order to accomplish this objective, not only must a sensor be in constant communication with a planning and scheduling system to process tasking requests, there must be an underlying framework to provide useful data products, such as angles only measurements. Existing optical signal processing implementations such as the Optical Processing Architecture at Lincoln (OPAL) are capable of converting mission data collections to angles only observations, but may be difficult for many users to obtain, support, and customize for low-cost missions and demonstration programs. The Ground Optical Signal Processing Architecture (GOSPA) will ingest raw imagery and telemetry data from a space based electro optical sensor and perform a background removal process to remove anomalous pixels, interpolate over bad pixels, and dominant temporal noise. After background removal, the streak end points and target centroids are located using a corner detection algorithm developed by Air Force Research Laboratory. These identified streak locations are then fused with the corresponding spacecraft telemetry data to determine the Right Ascension and Declination measurements with respect to time. To demonstrate the performance of GOSPA, non-rate tracking collections against a satellite in Geosynchronous Orbit are simulated from a visible optical imaging sensor in a polar Low Earth Orbit. Stars, noise and bad pixels are added to the simulated images based on look angles and sensor parameters. These collections are run through the GOSPA framework to provide angles- only measurements to the Air Force Research Laboratory Constrained Admissible Region Multiple Hypothesis Filter (CAR-MHF) in which an Initial Orbit Determination is performed and compared to truth data.

Galileo Jupiter approach orbit determination

NASA Technical Reports Server (NTRS)

Miller, J. K.; Nicholson, F. T.

1984-01-01

Orbit determination characteristics of the Jupiter approach phase of the Galileo mission are described. Predicted orbit determination performance is given for the various mission events that occur during Jupiter approach. These mission events include delivery of an atmospheric entry probe, acquisition of probe science data by the Galileo orbiter for relay to earth, delivery of an orbiter to a close encounter of the Galilean satellite Io, and insertion of the orbiter into orbit about Jupiter. The orbit determination strategy and resulting accuracies are discussed for the data types which include Doppler, range, optical imaging of Io, and a new Very Long Baseline Interferometry (VLBI) data type called Differential One-Way Range (DOR).

The Application of GIM in Precise Orbit Determination for LEO Satellites with Single-frequency GPS Measurements

NASA Astrophysics Data System (ADS)

Peng, D. J.; Wu, B.

2012-01-01

With the availability of precise GPS ephemeris and clock solution, the ionospheric range delay is left as the dominant error sources in the post-processing of space-borne GPS data from single-frequency receivers. Thus, the removal of ionospheric effects is a major prerequisite for an improved orbit reconstruction of LEO satellites equipped with low cost single-frequency GPS receivers. In this paper, the use of Global Ionospheric Maps (GIM) in kinematic and dynamic orbit determination for LEO satellites with single-frequency GPS measurements is discussed first,and then, estimating the scale factor of ionosphere to remove the ionospheric effects in C/A code pseudo-range measurements in both kinematic and adynamia orbit defemination approaches is addressed. As it is known the ionospheric path delay of space-borne GPS signals is strongly dependent on the orbit altitudes of LEO satellites, we selected real space-borne GPS data from CHAMP, GRACE, TerraSAR-X and SAC-C satellites with altitudes between 300 km and 800 km as sample data in this paper. It is demonstrated that the approach of eliminating ionospheric effects in space-borne C/A code pseudo-range by estimating the scale factor of ionosphere is highly effective. Employing this approach, the accuracy of both kinematic and dynamic orbits can be improved notably. Among those five LEO satellites, CHAMP with the lowest orbit altitude has the most remarkable orbit accuracy improvements, which are 55.6% and 47.6% for kinematic and dynamic approaches, respectively. SAC-C with the highest orbit altitude has the least orbit accuracy improvements accordingly, which are 47.8% and 38.2%, respectively.

Precise Orbital and Geodetic Parameter Estimation using SLR Observations for ILRS AAC

NASA Astrophysics Data System (ADS)

Kim, Young-Rok; Park, Eunseo; Oh, Hyungjik Jay; Park, Sang-Young; Lim, Hyung-Chul; Park, Chandeok

2013-12-01

In this study, we present results of precise orbital geodetic parameter estimation using satellite laser ranging (SLR) observations for the International Laser Ranging Service (ILRS) associate analysis center (AAC). Using normal point observations of LAGEOS-1, LAGEOS-2, ETALON-1, and ETALON-2 in SLR consolidated laser ranging data format, the NASA/ GSFC GEODYN II and SOLVE software programs were utilized for precise orbit determination (POD) and finding solutions of a terrestrial reference frame (TRF) and Earth orientation parameters (EOPs). For POD, a weekly-based orbit determination strategy was employed to process SLR observations taken from 20 weeks in 2013. For solutions of TRF and EOPs, loosely constrained scheme was used to integrate POD results of four geodetic SLR satellites. The coordinates of 11 ILRS core sites were determined and daily polar motion and polar motion rates were estimated. The root mean square (RMS) value of post-fit residuals was used for orbit quality assessment, and both the stability of TRF and the precision of EOPs by external comparison were analyzed for verification of our solutions. Results of post-fit residuals show that the RMS of the orbits of LAGEOS-1 and LAGEOS-2 are 1.20 and 1.12 cm, and those of ETALON-1 and ETALON-2 are 1.02 and 1.11 cm, respectively. The stability analysis of TRF shows that the mean value of 3D stability of the coordinates of 11 ILRS core sites is 7.0 mm. An external comparison, with respect to International Earth rotation and Reference systems Service (IERS) 08 C04 results, shows that standard deviations of polar motion XP and YP are 0.754 milliarcseconds (mas) and 0.576 mas, respectively. Our results of precise orbital and geodetic parameter estimation are reasonable and help advance research at ILRS AAC.

Evaluation of electrostatic charge effects on the data processing system and the orbiter communication and tracking receivers

NASA Technical Reports Server (NTRS)

Lawton, R. M.

1975-01-01

An analysis of radiated interference test results obtained from frictionally charged Orbiter TPS tile was presented. The tests included the measurement of noise pick-up by Orbiter S-band, L-band, C-band, and Ku-band antennas located beneath the tiles in a manner simulating their installation on Orbiter. In addition, the radiated field characteristics resulting from the static discharge was determined. The results are analyzed as to their effect on data bus equipment and on Orbiter Communications and Tracking (C&T) receivers. It was concluded that the radiated interference should have no effect on MDM's. However the CPU, IOP and PMU enclosures require some minor modification to assure immunity from P-static interference. Orbiter antenna tests indicate that the S-band receiver should not be affected by P-static noise. The TACAN and Radar Altimeter performance appears to be adequate but with a small margin. MSBLS performance is uncertain because laboratory instrumentation cannot approach the MSBLS sensitivity.

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.

High accuracy GNSS based navigation in GEO

NASA Astrophysics Data System (ADS)

Capuano, Vincenzo; Shehaj, Endrit; Blunt, Paul; Botteron, Cyril; Farine, Pierre-André

2017-07-01

Although significant improvements in efficiency and performance of communication satellites have been achieved in the past decades, it is expected that the demand for new platforms in Geostationary Orbit (GEO) and for the On-Orbit Servicing (OOS) on the existing ones will continue to rise. Indeed, the GEO orbit is used for many applications including direct broadcast as well as communications. At the same time, Global Navigation Satellites System (GNSS), originally designed for land, maritime and air applications, has been successfully used as navigation system in Low Earth Orbit (LEO) and its further utilization for navigation of geosynchronous satellites becomes a viable alternative offering many advantages over present ground based methods. Following our previous studies of GNSS signal characteristics in Medium Earth Orbit (MEO), GEO and beyond, in this research we specifically investigate the processing of different GNSS signals, with the goal to determine the best navigation performance they can provide in a GEO mission. Firstly, a detailed selection among different GNSS signals and different combinations of them is discussed, taking into consideration the L1 and L5 frequency bands, and the GPS and Galileo constellations. Then, the implementation of an Orbital Filter is summarized, which adaptively fuses the GN1SS observations with an accurate orbital forces model. Finally, simulation tests of the navigation performance achievable by processing the selected combination of GNSS signals are carried out. The results obtained show an achievable positioning accuracy of less than one meter. In addition, hardware-in-the-loop tests are presented using a COTS receiver connected to our GNSS Spirent simulator, in order to collect real-time hardware-in-the-loop observations and process them by the proposed navigation module.

On the Determination of the Orbits of Comets

NASA Astrophysics Data System (ADS)

Englefield, Henry

2013-06-01

Preface; 1. General view of the method; 2. On the motion of the point of intersection of the radius vector and cord; 3. On the comparison of the parabolic cord with the space which answers to the mean velocity of the earth in the same time; 4. Of the reduction of the second longitude of the comet; 5. On the proportion of the three curtate distances of the comet from the earth; 6. Of the graphical declination of the orbit of the earth; 7. Of the numerical quantities to be prepared for the construction or computation of the comet's orbit; 8. Determination of the distances of the comet from the earth and the sun; 9. Determination of the elements of the orbit from the determined distances; 10. Determination of the place of the comet from the earth and sun; 11. Determination of the distances of the comet from the earth and sun; 12. Determination of the comet's orbit; 13. Determination of the place of the comet; 14. Application of the graphical method to the comet of 1769; 15. Application of the distances found; 16. Determination of the place of the comet, for another given time; 17. Application of the trigonometrical method to the comet of 1769; 18. Determination of the elements of the orbit of the comet of 1769; Example of the graphical operation for the orbit of the comet of 1769; Example of the trigonometrical operation for the orbit of the comet of 1769; Conclusion; La Place's general method for determining the orbits of comets; Determination of the two elements of the orbit; Application of La Place's method of finding the approximate perihelion distance; Application of La Place's method for correcting the orbit of a comet, to the comet of 1769; Explanation and use of the tables; Tables; Appendix; Plates.

Researches on the Orbit Determination and Positioning of the Chinese Lunar Exploration Program

NASA Astrophysics Data System (ADS)

Li, P. J.

2015-07-01

This dissertation studies the precise orbit determination (POD) and positioning of the Chinese lunar exploration spacecraft, emphasizing the variety of VLBI (very long baseline interferometry) technologies applied for the deep-space exploration, and their contributions to the methods and accuracies of the precise orbit determination and positioning. In summary, the main contents are as following: In this work, using the real-time data measured by the CE-2 (Chang'E-2) detector, the accuracy of orbit determination is analyzed for the domestic lunar probe under the present condition, and the role played by the VLBI tracking data is particularly reassessed through the precision orbit determination experiments for CE-2. The experiments of the short-arc orbit determination for the lunar probe show that the combination of the ranging and VLBI data with the arc of 15 minutes is able to improve the accuracy by 1-1.5 order of magnitude, compared to the cases for only using the ranging data with the arc of 3 hours. The orbital accuracy is assessed through the orbital overlapping analysis, and the results show that the VLBI data is able to contribute to the CE-2's long-arc POD especially in the along-track and orbital normal directions. For the CE-2's 100 km× 100 km lunar orbit, the position errors are better than 30 meters, and for the CE-2's 15 km× 100 km orbit, the position errors are better than 45 meters. The observational data with the delta differential one-way ranging (Δ DOR) from the CE-2's X-band monitoring and control system experimental are analyzed. It is concluded that the accuracy of Δ DOR delay is dramatically improved with the noise level better than 0.1 ns, and the systematic errors are well calibrated. Although it is unable to support the development of an independent lunar gravity model, the tracking data of CE-2 provided the evaluations of different lunar gravity models through POD, and the accuracies are examined in terms of orbit-to-orbit solution differences for several gravity models. It is found that for the 100 km× 100 km lunar orbit, with a degree and order expansion up to 165, the JPL's gravity model LP165P does not show noticeable improvement over Japan's SGM series models (100× 100), but for the 15 km× 100 km lunar orbit, a higher degree-order model can significantly improve the orbit accuracy. After accomplished its nominal mission, CE-2 launched its extended missions, which involving the L2 mission and the 4179 Toutatis mission. During the flight of the extended missions, the regime offers very little dynamics thus requires an extensive amount of time and tracking data in order to attain a solution. The overlap errors are computed, and it is indicated that the use of VLBI measurements is able to increase the accuracy and reduce the total amount of tracking time. An orbit determination method based on the polynomial fitting is proposed for the CE-3's planned lunar soft landing mission. In this method, spacecraft's dynamic modeling is not necessary, and its noise reduction is expected to be better than that of the point positioning method by making full use of all-arc observational data. The simulation experiments and real data processing showed that the optimal description of the CE-1's free-fall landing trajectory is a set of five-order polynomial functions for each of the position components as well as velocity components in J2000.0. The combination of the VLBI delay, the delay rate data, and the USB (united S-band) ranging data significantly improved the accuracy than the use of USB data alone. In order to determine the position for the CE-3's Lunar Lander, a kinematic statistical method is proposed. This method uses both ranging and VLBI measurements to the lander for a continuous arc, combing with precise knowledge about the motion of the moon as provided by planetary ephemeris, to estimate the lander's position on the lunar surface with high accuracy. Application of the lunar digital elevation model (DEM) as constraints in the lander positioning is helpful. The positioning method for the traverse of lunar rover is also investigated. The integration of delay-rate method is able to achieve higher precise positioning results than the point positioning method. This method provides a wide application of the VLBI data. In the automated sample return mission, the lunar orbit rendezvous and docking are involved. Precise orbit determination using the same-beam VLBI (SBI) measurement for two spacecraft at the same time is analyzed. The simulation results showed that the SBI data is able to improve the absolute and relative orbit accuracy for two targets by 1-2 orders of magnitude. In order to verify the simulation results and test the two-target POD software developed by SHAO (Shanghai Astronomical Observatory), the real SBI data of the SELENE (Selenological and Engineering Explorer) are processed. The POD results for the Rstar and the Vstar showed that the combination of SBI data could significantly improve the accuracy for the two spacecraft, especially for the Vstar with less ranging data, and the POD accuracy is improved by approximate one order of magnitude to the POD accuracy of the Rstar.

Single-Frequency GPS Relative Navigation in a High Ionosphere Orbital Environment

NASA Technical Reports Server (NTRS)

Conrad, Patrick R.; Naasz, Bo J.

2007-01-01

The Global Positioning System (GPS) provides a convenient source for space vehicle relative navigation measurements, especially for low Earth orbit formation flying and autonomous rendezvous mission concepts. For single-frequency GPS receivers, ionospheric path delay can be a significant error source if not properly mitigated. In particular, ionospheric effects are known to cause significant radial position error bias and add dramatically to relative state estimation error if the onboard navigation software does not force the use of measurements from common or shared GPS space vehicles. Results from GPS navigation simulations are presented for a pair of space vehicles flying in formation and using GPS pseudorange measurements to perform absolute and relative orbit determination. With careful measurement selection techniques relative state estimation accuracy to less than 20 cm with standard GPS pseudorange processing and less than 10 cm with single-differenced pseudorange processing is shown.

Some significant considerations in the planning of sortie missions. [of space transportation system

NASA Technical Reports Server (NTRS)

Loftus, J. P., Jr.; Cour-Palais, B. G.; Moore, J. W.; Lohman, R. L.

1980-01-01

Opportunities and limitations to be considered in the planning of Space Shuttle/Spacelab sortie missions are discussed. As shown by a simple model of the flow of STS equipment through ground processing and flights under ideal conditions, mission duration is constrained by Orbiter availability, which is determined initially by the Orbiter production schedule and the turnaround time required between missions, and by the usage rate and quantity limitations of mission consumables. Additional considerations affecting mission duration include reductions in crew productivity upon increased mission duration and crew size, spacecraft and experiment degradation, equipment and processing facility cost effectiveness, and requirements for a power extension package, which considerations imply that increased allowable landing weight would make co-manifesting (the combination of Spacelab and deliverable payload missions) more attractive. Advantages related to payload recoverability, human presence, ease of access and the availability of different orbits are also pointed out.

Satellite orbit considerations for a global change technology architecture trade study

NASA Technical Reports Server (NTRS)

Harrison, Edwin F.; Gibson, Gary G.; Suttles, John T.; Buglia, James J.; Taback, Israel

1991-01-01

A study was conducted to determine satellite orbits for earth observation missions aimed at obtaining data for assessing data global climate change. A multisatellite system is required to meet the scientific requirements for temporal coverage over the globe. The best system consists of four sun-synchronous satellites equally spaced in local time of equatorial crossing. This system can obtain data every three hours for all regions. Several other satellite systems consisting of combinations of sun-synchronous orbits and either the Space Station Freedom or a mid-altitude equatorial satellite can provide three to six hour temporal coverage, which is sufficient for measuring many of the parameters required for the global change monitoring mission. Geosynchronous satellites are required to study atmospheric and surface processes involving variations on the order of a few minutes to an hour. One or two geosynchronous satellites can be relocated in longitude to study processes over selected regions of earth.

Development of the GPM Observatory Thermal Vacuum Test Model

NASA Technical Reports Server (NTRS)

Yang, Kan; Peabody, Hume

2012-01-01

A software-based thermal modeling process was documented for generating the thermal panel settings necessary to simulate worst-case on-orbit flight environments in an observatory-level thermal vacuum test setup. The method for creating such a thermal model involved four major steps: (1) determining the major thermal zones for test as indicated by the major dissipating components on the spacecraft, then mapping the major heat flows between these components; (2) finding the flight equivalent sink temperatures for these test thermal zones; (3) determining the thermal test ground support equipment (GSE) design and initial thermal panel settings based on the equivalent sink temperatures; and (4) adjusting the panel settings in the test model to match heat flows and temperatures with the flight model. The observatory test thermal model developed from this process allows quick predictions of the performance of the thermal vacuum test design. In this work, the method described above was applied to the Global Precipitation Measurement (GPM) core observatory spacecraft, a joint project between NASA and the Japanese Aerospace Exploration Agency (JAXA) which is currently being integrated at NASA Goddard Space Flight Center for launch in Early 2014. From preliminary results, the thermal test model generated from this process shows that the heat flows and temperatures match fairly well with the flight thermal model, indicating that the test model can simulate fairly accurately the conditions on-orbit. However, further analysis is needed to determine the best test configuration possible to validate the GPM thermal design before the start of environmental testing later this year. Also, while this analysis method has been applied solely to GPM, it should be emphasized that the same process can be applied to any mission to develop an effective test setup and panel settings which accurately simulate on-orbit thermal environments.

Automatic Reacquisition of Satellite Positions by Detecting Their Expected Streaks in Astronomical Images

NASA Astrophysics Data System (ADS)

Levesque, M.

Artificial satellites, and particularly space junk, drift continuously from their known orbits. In the surveillance-of-space context, they must be observed frequently to ensure that the corresponding orbital parameter database entries are up-to-date. Autonomous ground-based optical systems are periodically tasked to observe these objects, calculate the difference between their predicted and real positions and update object orbital parameters. The real satellite positions are provided by the detection of the satellite streaks in the astronomical images specifically acquired for this purpose. This paper presents the image processing techniques used to detect and extract the satellite positions. The methodology includes several processing steps including: image background estimation and removal, star detection and removal, an iterative matched filter for streak detection, and finally false alarm rejection algorithms. This detection methodology is able to detect very faint objects. Simulated data were used to evaluate the methodology's performance and determine the sensitivity limits where the algorithm can perform detection without false alarm, which is essential to avoid corruption of the orbital parameter database.

Orbit and clock determination of BDS regional navigation satellite system based on IGS M-GEX and WHU BETS tracking network

NASA Astrophysics Data System (ADS)

GENG, T.; Zhao, Q.; Shi, C.; Shum, C.; Guo, J.; Su, X.

2013-12-01

BeiDou Navigation Satellite System (BDS) began to provide the regional open service on December 27th 2012 and will provide the global open service by the end of 2020. Compared to GPS, the space segment of BDS Regional System consists of 5 Geostationary Earth Orbit satellites (GEO), 5 Inclined Geosynchronous Orbit satellites (IGSO) and 4 Medium Earth orbit (MEO) satellites. Since 2011, IGS Multiple-GNSS Experiment (M-GEX) focuses on tracking the newly available GNSS signals. This includes all signals from the modernized satellites of the GPS and GLONASS systems, as well as signals of the BDS, Galileo and QZSS systems. Up to now, BDS satellites are tracked by around 25 stations with a variety of different antennas and receivers from different GNSS manufacture communities in M-GEX network. Meanwhile, there are 17 stations with Unicore Communications Incorporation's GPS/BDS receivers in BeiDou Experimental Tracking Stations (BETS) network by Wuhan University. In addition, 5 BDS satellites have been tracking by the International Laser Ranging Service (ILRS). BDS performance is expected to be further studied by the GNSS communities. Following an introduction of the BDS system and above different tracking network, this paper discusses the achieved BDS characterization and performance assessment. Firstly, the BDS signal and measurement quality are analyzed with different antennas and receivers in detail compared to GPS. This includes depth of coverage for satellite observation, carrier-to-noise-density ratios, code noise and multipath, carrier phase errors. Secondly, BDS Precise Orbit Determination (POD) is processed. Different arc lengths and sets of orbit parameters are tested using Position And Navigation Data Analysis software (PANDA) which is developed at the Wuhan University. GEO, IGSO and MEO satellites orbit quality will be assessed using overlap comparison, 2-day orbit fit and external validations with Satellite Laser Range (SLR). Then BDS satellites are equipped with Rubidium clocks and clocks performance are also presented. Finally, benefits of BDS processing strategies and further developments are concluded.

Ground Optical Signal Processing Architecture for Contributing Space-Based SSA Sensor Data

DTIC Science & Technology

2014-09-01

Where 2 zodiacal 2 thermalphotons dNdNsignaldN  and readdN is the read noise in noise-electrons. dNthermal is the photoelectron noise due...PhD, USAF Defense Advanced Research Projects Agency, Arlington, VA. ABSTRACT DARPA’s OrbitOutlook aims to augment the performance of the Space...SDA) and determine when satellites are at risk. OrbitOutlook also seeks to demonstrate the ability to rapidly include new instruments to alert for

Shuttle/payload communications and data systems interface analysis

NASA Technical Reports Server (NTRS)

Huth, G. K.

1980-01-01

The payload/orbiter functional command signal flow and telemetry signal flow are discussed. Functional descriptions of the various orbiter communication/avionic equipment involved in processing a command to a payload either from the ground through the orbiter by the payload specialist on the orbiter are included. Functional descriptions of the various orbiter communication/avionic equipment involved in processing telemetry data by the orbiter and transmitting the processed data to the ground are presented. The results of the attached payload/orbiter single processing and data handling system evaluation are described. The causes of the majority of attached payload/orbiter interface problems are delineated. A refined set of required flux density values for a detached payload to communicate with the orbiter is presented.

Thermospheric density estimation from SLR observations of LEO satellites - A case study with the ANDE-Pollux satellite

NASA Astrophysics Data System (ADS)

Blossfeld, M.; Schmidt, M.; Erdogan, E.

2016-12-01

The thermospheric neutral density plays a crucial role within the equation of motion of Earth orbiting objects since drag, lift or side forces are one of the largest non-gravitational perturbations acting on the satellite. Precise Orbit Determination (POD) methods can be used to estimate thermospheric density variations from measured orbit determinations. One method which provides highly accurate measurements of the satellite position is Satellite Laser Ranging (SLR). Within the POD process, scaling factors are estimated frequently. These scaling factors can be either used for the scaling of the so called satellite-specific drag (ballistic) coefficients or the integrated thermospheric neutral density. We present a method for analytically model the drag coefficient based on a couple of physical assumptions and key parameters. In this paper, we investigate the possibility to use SLR observations to the very low Earth orbiting satellite ANDE-Pollux (approximately at 350km altitude) to determine scaling factors for different a priori thermospheric density models. We perform a POD for ANDE-Pollux covering 49 days between August 2009 and September 2009 which means the time span containing the largest number of observations during the short lifetime of the satellite. Finally, we compare the obtained scaled thermospheric densities w.r.t. each other

Cassini's Maneuver Automation Software (MAS) Process: How to Successfully Command 200 Navigation Maneuvers

NASA Technical Reports Server (NTRS)

Yang, Genevie Velarde; Mohr, David; Kirby, Charles E.

2008-01-01

To keep Cassini on its complex trajectory, more than 200 orbit trim maneuvers (OTMs) have been planned from July 2004 to July 2010. With only a few days between many of these OTMs, the operations process of planning and executing the necessary commands had to be automated. The resulting Maneuver Automation Software (MAS) process minimizes the workforce required for, and maximizes the efficiency of, the maneuver design and uplink activities. The MAS process is a well-organized and logically constructed interface between Cassini's Navigation (NAV), Spacecraft Operations (SCO), and Ground Software teams. Upon delivery of an orbit determination (OD) from NAV, the MAS process can generate a maneuver design and all related uplink and verification products within 30 minutes. To date, all 112 OTMs executed by the Cassini spacecraft have been successful. MAS was even used to successfully design and execute a maneuver while the spacecraft was in safe mode.

GPS Navigation Results from the Low Power Transceiver CANDOS Experiment on STS-107

NASA Technical Reports Server (NTRS)

Haas, Lin; Massey, Chris; Baraban, Dmitri; Kelbel, David; Lee, Taesul; Long, Anne; Carpenter, J. Russell

2003-01-01

This paper presents the Global Positioning System (GPS) navigation results from the Communications and Savigation Demonstration on Shuttle (CANDOS) experiment flown on STS- 107. The CAkDOS experiment consisted of the Low Power Transceiver (LPT) that hosted the GPS Enhanced Orbit Determination Experiment (GEODE) orbit determination software. All CANDOS test data were recovered during the mission using the LPT's Tracking and Data Relay Satellite System (TDRSS) uplinh'downlink communications capabilit! . An overview of the LPT's navigation software and the GPS experiment timeline is presented. In addition. this paper discusses GEODE performance results. including comparisons ibith the Best Estimate of Trajectory (BET). N.ASA Johnson Space Center (JSC) real-time ground navigation vectors. and post-processed solutions using the Goddard Trajectory Determination System (GTDS).

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…

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.

A review of cooperative and uncooperative spacecraft pose determination techniques for close-proximity operations

NASA Astrophysics Data System (ADS)

Opromolla, Roberto; Fasano, Giancarmine; Rufino, Giancarlo; Grassi, Michele

2017-08-01

The capability of an active spacecraft to accurately estimate its relative position and attitude (pose) with respect to an active/inactive, artificial/natural space object (target) orbiting in close-proximity is required to carry out various activities like formation flying, on-orbit servicing, active debris removal, and space exploration. According to the specific mission scenario, the pose determination task involves both theoretical and technological challenges related to the search for the most suitable algorithmic solution and sensor architecture, respectively. As regards the latter aspect, electro-optical sensors represent the best option as their use is compatible with mass and power limitation of micro and small satellites, and their measurements can be processed to estimate all the pose parameters. Overall, the degree of complexity of the challenges related to pose determination largely varies depending on the nature of the targets, which may be actively/passively cooperative, uncooperative but known, or uncooperative and unknown space objects. In this respect, while cooperative pose determination has been successfully demonstrated in orbit, the uncooperative case is still under study by universities, research centers, space agencies and private companies. However, in both the cases, the demand for space applications involving relative navigation maneuvers, also in close-proximity, for which pose determination capabilities are mandatory, is significantly increasing. In this framework, a review of state-of-the-art techniques and algorithms developed in the last decades for cooperative and uncooperative pose determination by processing data provided by electro-optical sensors is herein presented. Specifically, their main advantages and drawbacks in terms of achieved performance, computational complexity, and sensitivity to variability of pose and target geometry, are highlighted.

SHINE, The SpHere INfrared survey for Exoplanets

NASA Astrophysics Data System (ADS)

Chauvin, G.; Desidera, S.; Lagrange, A.-M.; Vigan, A.; Feldt, M.; Gratton, R.; Langlois, M.; Cheetham, A.; Bonnefoy, M.; Meyer, M.

2017-12-01

The SHINE survey for SPHERE High-contrast ImagiNg survey for Exoplanets, is a large near-infrared survey of 400-600 young, nearby stars and represents a significant component of the SPHERE consortium Guaranteed Time Observations consisting in 200 observing nights. The scientific goals are: i) to characterize known planetary systems (architecture, orbit, stability, luminosity, atmosphere); ii) to search for new planetary systems using SPHERE's unprecedented performance; and finally iii) to determine the occurrence and orbital and mass function properties of the wide-orbit, giant planet population as a function of the stellar host mass and age. Combined, the results will increase our understanding of planetary atmospheric physics and the processes of planetary formation and evolution.

Minor Planet Center Data Processing Challenges

NASA Astrophysics Data System (ADS)

Rudenko, Michael

2015-08-01

The Minor Planet Center (MPC) is the single worldwide location for receipt and distribution of positional measurements of minor planets, comets and outer irregular natural satellites of the major planets. The MPC is responsible for the identification, designation and orbit computation for all of these objects.Over 2 million observations are received each month via the internet, and are validated and processed in near real time. The observations come in batches whose formats are checked and whose observations are run through a number of other routine checks such as departure from great circle motion, prior publication, single observations, near duplicates, etc. Some or all of a batch of observations may be returned to its sender if they fail one or more of the checks. After the observations have been validated, they are processed to produce orbits of newly discovered objects or used to update the orbits of known objects.Given the volume of observations, the sheer number of known objects against which to possibly match, the shortness of the time interval over which each object was likely observed, and the uncertainties in the positions, and occasionally possible errors in times, reported, a number of data processing challenges face the MPC. These include the following: Identifying observations of objects reported as new with already known objects; linking together sets of observations from different nights (possibly at different apparitions) which may belong to the same object; determining if a set of observations has been assigned to the wrong object; determining if an object with a very short arc is possibly a Near-Earth object; determining and examining the range of possible variant orbits of newly discovered Near-Earth objects with very short observation arcs for cases which indicate an object is potentially on a collision course with Earth; linking observations to known artificial satellites and/or booster stages and other space "junk"; prioritizing newly discovered objects in order of need of follow up; and, efficiently matching one or more observations with known objects (MPChecker web service), possibly of observations in the distant past.

Calculating wave-generated bottom orbital velocities from surface-wave parameters

USGS Publications Warehouse

Wiberg, P.L.; Sherwood, C.R.

2008-01-01

Near-bed wave orbital velocities and shear stresses are important parameters in many sediment-transport and hydrodynamic models of the coastal ocean, estuaries, and lakes. Simple methods for estimating bottom orbital velocities from surface-wave statistics such as significant wave height and peak period often are inaccurate except in very shallow water. This paper briefly reviews approaches for estimating wave-generated bottom orbital velocities from near-bed velocity data, surface-wave spectra, and surface-wave parameters; MATLAB code for each approach is provided. Aspects of this problem have been discussed elsewhere. We add to this work by providing a method for using a general form of the parametric surface-wave spectrum to estimate bottom orbital velocity from significant wave height and peak period, investigating effects of spectral shape on bottom orbital velocity, comparing methods for calculating bottom orbital velocity against values determined from near-bed velocity measurements at two sites on the US east and west coasts, and considering the optimal representation of bottom orbital velocity for calculations of near-bed processes. Bottom orbital velocities calculated using near-bed velocity data, measured wave spectra, and parametric spectra for a site on the northern California shelf and one in the mid-Atlantic Bight compare quite well and are relatively insensitive to spectral shape except when bimodal waves are present with maximum energy at the higher-frequency peak. These conditions, which are most likely to occur at times when bottom orbital velocities are small, can be identified with our method as cases where the measured wave statistics are inconsistent with Donelan's modified form of the Joint North Sea Wave Project (JONSWAP) spectrum. We define the 'effective' forcing for wave-driven, near-bed processes as the product of the magnitude of forcing times its probability of occurrence, and conclude that different bottom orbital velocity statistics may be appropriate for different problems. ?? 2008 Elsevier Ltd.

Copernicus POD Service Operations

NASA Astrophysics Data System (ADS)

Fernandez, Jaime; Ayuga, Francisco; Fernandez, Carlos; Peter, Heike; Femenias, Pierre

2016-08-01

The Copernicus POD (Precise Orbit Determination) Serviceis part of the Copernicus PDGS Ground Segment of the Sentinel missions. A GMV-led consortium is operating the Copernicus POD Service being in charge of generating precise orbital products and auxiliary data files of Sentinel-1, -2, & -3 missions, for their use as part of the processing chains of the respective Sentinel PDGS.Although the characteristics and the requirements are different for the three missions, the same core POD setup is used to the largest possible extent. At the moment, the CPOD Service is operating 3 satellites: Sentinel-1A, -2A and -3A, and is ready to support operations of Sentinel-1B.This paper presents the status of the CPOD Service in terms of operations and orbital accuracy achieved for the different orbit products of the different missions, focusing on Sentinel-3A preliminary results.

DPOD2014: a new DORIS extension of ITRF2014 for Precise Orbit Determination

NASA Astrophysics Data System (ADS)

Moreaux, G.; Willis, P.; Lemoine, F. G.; Zelensky, N. P.

2016-12-01

As one of the tracking systems used to determine orbits of the altimeter mission satellites (such as TOPEX/Poseidon, Envisat, Jason-1/2/3 & Cryosat-2), the position of the DORIS tracking stations provides a fundamental reference for the estimation of the precise orbits and so, by extension is fundamental for the quality of the altimeter data and derived products. Therefore, the time evolution of the position of both the existing and the newest DORIS stations must be precisely modeled and regularly updated. To satisfy operational requirements for precise orbit determination and routine delivery of geodetic products, the International DORIS Service maintains the so-called DPOD solutions, which can be seen as extensions of the latest available ITRF solution from the International Earth Rotation and Reference Systems Service (IERS). In mid-2016, the IDS agreed to change the processing strategy of the DPOD solution. The new solution from the IDS Combination Center (CC) consists of a DORIS cumulative position and velocity solution using the latest IDS combined weekly solutions. The first objective of this study is to describe the new DPOD elaboration scheme and to show the IDS CC internal validation steps. The second purpose is to present the external validation process made by an external team before the new DPOD is made available to all the users. The elaboration and validation procedures will be illustrated by the presentation of first version of the DPOD2014 (ITRF2014 DORIS extension) and focus will be given on the update of the position and velocity of two DORIS sites: Everest (after Gorkha earthquake M7.8 in April 2015) and Thule (Greenland).

Navigation strategy and filter design for solar electric missions

NASA Technical Reports Server (NTRS)

Tapley, B. D.; Hagar, H., Jr.

1972-01-01

Methods which have been proposed to improve the navigation accuracy for the low-thrust space vehicle include modifications to the standard Sequential- and Batch-type orbit determination procedures and the use of inertial measuring units (IMU) which measures directly the acceleration applied to the vehicle. The navigation accuracy obtained using one of the more promising modifications to the orbit determination procedures is compared with a combined IMU-Standard. The unknown accelerations are approximated as both first-order and second-order Gauss-Markov processes. The comparison is based on numerical results obtained in a study of the navigation requirements of a numerically simulated 152-day low-thrust mission to the asteroid Eros. The results obtained in the simulation indicate that the DMC algorithm will yield a significant improvement over the navigation accuracies achieved with previous estimation algorithms. In addition, the DMC algorithms will yield better navigation accuracies than the IMU-Standard Orbit Determination algorithm, except for extremely precise IMU measurements, i.e., gyroplatform alignment .01 deg and accelerometer signal-to-noise ratio .07. Unless these accuracies are achieved, the IMU navigation accuracies are generally unacceptable.

Orbit determination and prediction of GEO satellite of BeiDou during repositioning maneuver

NASA Astrophysics Data System (ADS)

Cao, Fen; Yang, XuHai; Li, ZhiGang; Sun, BaoQi; Kong, Yao; Chen, Liang; Feng, Chugang

2014-11-01

In order to establish a continuous GEO satellite orbit during repositioning maneuvers, a suitable maneuver force model has been established associated with an optimal orbit determination method and strategy. A continuous increasing acceleration is established by constructing a constant force that is equivalent to the pulse force, with the mass of the satellite decreasing throughout maneuver. This acceleration can be added to other accelerations, such as solar radiation, to obtain the continuous acceleration of the satellite. The orbit determination method and strategy are illuminated, with subsequent assessment of the orbit being determined and predicted accordingly. The orbit of the GEO satellite during repositioning maneuver can be determined and predicted by using C-Band pseudo-range observations of the BeiDou GEO satellite with COSPAR ID 2010-001A in 2011 and 2012. The results indicate that observations before maneuver do affect orbit determination and prediction, and should therefore be selected appropriately. A more precise orbit and prediction can be obtained compared to common short arc methods when observations starting 1 day prior the maneuver and 2 h after the maneuver are adopted in POD (Precise Orbit Determination). The achieved URE (User Range Error) under non-consideration of satellite clock errors is better than 2 m within the first 2 h after maneuver, and less than 3 m for further 2 h of orbit prediction.

Precise attitude determination of defunct satellite laser ranging tragets

NASA Astrophysics Data System (ADS)

Pittet, Jean-Noel; Schildknecht, Thomas; Silha, Jiri

2016-07-01

The Satellite Laser Ranging (SLR) technology is used to determine the dynamics of objects equipped with so-called retro-reflectors or retro-reflector arrays (RRA). This type of measurement allows to range to the spacecraft with very high precision, which leads to determination of very accurate orbits. Non-active spacecraft, which are not any more attitude controlled, tend to start to spin or tumble under influence of the external and internal torques. Such a spinning can be around one constant axis of rotation or it can be more complex, when also precession and nutation motions are present. The rotation of the RRA around the spacecraft's centre of mass can create both a oscillation pattern of laser range signal and a periodic signal interruption when the RRA is hidden behind the satellite. In our work we will demonstrate how the SLR ranging technique to cooperative targets can be used to determine precisely their attitude state. The processing of the obtained data will be discussed, as well as the attitude determination based on parameters estimation. Continuous SLR measurements to one target can allow to accurately monitor attitude change over time which can be further used for the future attitude modelling. We will show our solutions of the attitude states determined for the non-active ESA satellite ENVISAT based on measurements acquired during year 2013-2015 by Zimmerwald SLR station, Switzerland. The angular momentum shows a stable behaviour with respect to the orbital plane but is not aligned with orbital momentum. The determination of the inertial rotation over time, shows it evolving between 130 to 190 seconds within two year. Parameter estimation also bring a strong indication of a retrograde rotation. Results on other former satellites in low and medium Earth orbit such as TOPEX/Poseidon or GLONASS type will be also presented.

An advanced analysis method of initial orbit determination with too short arc data

NASA Astrophysics Data System (ADS)

Li, Binzhe; Fang, Li

2018-02-01

This paper studies the initial orbit determination (IOD) based on space-based angle measurement. Commonly, these space-based observations have short durations. As a result, classical initial orbit determination algorithms give poor results, such as Laplace methods and Gauss methods. In this paper, an advanced analysis method of initial orbit determination is developed for space-based observations. The admissible region and triangulation are introduced in the method. Genetic algorithm is also used for adding some constraints of parameters. Simulation results show that the algorithm can successfully complete the initial orbit determination.

Trajectory optimization for an asymmetric launch vehicle. M.S. Thesis - MIT

NASA Technical Reports Server (NTRS)

Sullivan, Jeanne Marie

1990-01-01

A numerical optimization technique is used to fully automate the trajectory design process for an symmetric configuration of the proposed Advanced Launch System (ALS). The objective of the ALS trajectory design process is the maximization of the vehicle mass when it reaches the desired orbit. The trajectories used were based on a simple shape that could be described by a small set of parameters. The use of a simple trajectory model can significantly reduce the computation time required for trajectory optimization. A predictive simulation was developed to determine the on-orbit mass given an initial vehicle state, wind information, and a set of trajectory parameters. This simulation utilizes an idealized control system to speed computation by increasing the integration time step. The conjugate gradient method is used for the numerical optimization of on-orbit mass. The method requires only the evaluation of the on-orbit mass function using the predictive simulation, and the gradient of the on-orbit mass function with respect to the trajectory parameters. The gradient is approximated with finite differencing. Prelaunch trajectory designs were carried out using the optimization procedure. The predictive simulation is used in flight to redesign the trajectory to account for trajectory deviations produced by off-nominal conditions, e.g., stronger than expected head winds.

A Real-Time Orbit Determination Method for Smooth Transition from Optical Tracking to Laser Ranging of Debris

PubMed Central

Li, Bin; Sang, Jizhang; Zhang, Zhongping

2016-01-01

A critical requirement to achieve high efficiency of debris laser tracking is to have sufficiently accurate orbit predictions (OP) in both the pointing direction (better than 20 arc seconds) and distance from the tracking station to the debris objects, with the former more important than the latter because of the narrow laser beam. When the two line element (TLE) is used to provide the orbit predictions, the resultant pointing errors are usually on the order of tens to hundreds of arc seconds. In practice, therefore, angular observations of debris objects are first collected using an optical tracking sensor, and then used to guide the laser beam pointing to the objects. The manual guidance may cause interrupts to the laser tracking, and consequently loss of valuable laser tracking data. This paper presents a real-time orbit determination (OD) and prediction method to realize smooth and efficient debris laser tracking. The method uses TLE-computed positions and angles over a short-arc of less than 2 min as observations in an OD process where simplified force models are considered. After the OD convergence, the OP is performed from the last observation epoch to the end of the tracking pass. Simulation and real tracking data processing results show that the pointing prediction errors are usually less than 10″, and the distance errors less than 100 m, therefore, the prediction accuracy is sufficient for the blind laser tracking. PMID:27347958

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.

Performance Evaluation of Orbit Determination System during Initial Phase of INSAT-3 Mission

NASA Astrophysics Data System (ADS)

Subramanian, B.; Vighnesam, N. V.

INSAT-3C is the second in the third generation of ISRO's INSAT series of satellites that was launched by ARIANE-SPACE on 23 January 2002 at 23 h 46 m 57 s (lift off time in U.T). The ARIANE-4 Flight Nr.147 took off from Kourou in French Guyana and injected the 2750-kg communications satellite in a geostationary transfer orbit of (571 X 35935) km with an inclination of 4.007 deg at 00 h 07 m 48 s U.T on 24 January 2002 (1252 s after lift off). The satellite was successfully guided into its intended geostationary position of 74 deg E longitude by 09 February 2002 after a series of four firings of its Liquid Apogee Motor (LAM) and four station acquisition (STAQ) maneuvers. Six distinct phases of the mission were categorized based on the orbit characteristics of the INSAT- 3C mission, namely, the pre-launch phase, the launch phase, transfer orbit phase, intermediate orbit phase, drift orbit phase and synchronous orbit phase. The orbit with a perigee height of 571 km at injection of the satellite, was gradually raised to higher orbits with perigee height increasing to 9346 km after Apogee Motor Firing #1 (AMF #1), 18335 km after AMF #2, 32448 km after AMF #3 and 35493 km after AMF #4. The North and South solar panels and the reflectors were deployed at this stage of the mission and the attitude of the satellite with respect to the three axes was stabilized. The Orbit Determination System (ODS) that was used in the initial phase of the mission played a crucial role in realizing the objectives of the mission. This system which consisted of Tracking Data Pre-Processing (TDPP) software, Ephemeris Generation (EPHGEN) software and the Orbit Determination (OD) software, performed rigorously and its results were used for planning the AMF and STAQ strategies with a greater degree of accuracy. This paper reports the results of evaluation of the performance of the apogee-motor firings employed to place the satellite in its intended position where it is collocated with INSAT-1D satellite. The orbit of the satellite had to be determined continuously at each stage of the initial phase of the mission at a brisk pace and this study shows that the ODS provided consistent results to meet the stringent requirements of the mission operations. At each stage of the mission the orbit was determined using tracking data obtained over varying periods of time. The orbit solutions obtained from short arc OD's are compared with that obtained using the longest arc OD of each stage of the initial phase of the mission. The results of this study have been tabled in this paper. The performance of the ODS in calibrating the ARIANE-4 launch vehicle has been analyzed. A comparison of the orbit elements obtained from the mission operational ODS with the injection parameters provided by CNES, Centre Spatial Guyanais has been made in this paper which shows that the satellite was injected well within the 1 dispersions quoted by ARIANE-SPACE. A comparison has also been shown between the determined transfer orbit elements with pre-launch nominal orbit elements. For the initial phase of this mission ranging support was provided by Hassan earth station at India and INMARSAT network of stations at LakeCowichan (Canada), Fucino (Italy) and Beijing (China). The performance of the tracking systems employed by these stations has been studied. The quality of tracking data obtained from these stations has also been assessed.

Supporting the Copernicus POD Service

NASA Astrophysics Data System (ADS)

Peter, Heike; Springer, Tim; Otten, Michiel; Fernandez, Jaime; Escobar, Diego; Femenias, Pierre

2015-12-01

The Copernicus POD (Precise Orbit Determination) Service is part of the Copernicus PDGS (Payload Data Ground Segment) of the Sentinel 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. As part of the consortium PosiTim is responsible for implementing and testing software and model updates thoroughly before integrating them in the operational chain of the Copernicus POD Service. The NAPEOS (Navigation Package for Earth Observation Satellites) software is used for the generation of the orbit products within the Copernicus POD Service. The test procedures and results obtained for a recent software and model update to IERS 2010 Conventions are presented. It has been tested as well that the arc length of 72 hours for the non-time critical (NTC) orbit solutions might be shorten to 48 hours without losing accuracy. Orbit comparisons to external solutions help to validate the different orbit solutions. GPS antenna phase centre variations (PCVs) are one of the largest systematic error sources in POD. Since the satellite body may cause signal multipath a ground calibration of the GPS antenna without taking into account the satellite body might not be sufficient to quantify the PCVs. The PCVs are therefore obtained by an in-flight calibration. A first map for the PCVs determined from a limited amount of data at the beginning of the mission has shown significant multipath signals in parts of the antenna for code and carrier phase measurements. Since the satellite has moving parts it has been checked carefully if these multipath regions are moving as well or if they are antenna-fixed. Normally the correction maps are only applied for the carrier phase measurements. Since significant multipath has been spotted for the code measurements as well investigations are performed to study the impact of additionally applying code correction maps in the POD process.

Cassini/Huygens Science Instruments, Spacecraft, and Mission

NASA Technical Reports Server (NTRS)

Jaffe, Leonard D.; Herrell, Linda M.

1997-01-01

The Cassini spacecraft will take 18 scientific instruments to Saturn. After launch and a seven-year cruise, Cassini will arrive at Saturn and separate into a Saturn orbiter and an atmospheric probe, called Huygens, which will descend to the surface of Titan. The orbiter will orbit the planet for four years, making close flybys of five satellites, including multiple flybys of Titan. Communication with Earth is at X-band; the maximum downlink rate from Saturn is 166 x 10(exp 3) bps. Orbiter instruments are body mounted; the spacecraft must be turned to point some of them toward objects of interest. The orbiter carries 12 instruments. Optical instruments provide imagery and spectrometry. Radar supplies imaging, altimetry, and radiometry. Radio links contribute information about intervening material and gravity fields. Other instruments measure electromagnetic fields and the properties of plasma, energetic particles, and dust particles. The probe is spin stabilized. It returns data via an S-band link to the orbiter. The probe's six instruments include sensors to determine atmospheric physical properties and composition. Radiometric and optical sensors will produce data on thermal balance and obtain images of Titan's atmosphere and surface. Doppler measurements between probe and orbiter will provide wind profiles. Surface sensors will measure impact acceleration, thermal and electrical properties, and, if the surface is liquid, density and refractive index. This design will enable Cassini to determine the composition; the physical, morphological, and geological nature; and the physical and chemical processes of the atmospheres, surfaces, and magnetosphere of the Saturnian system. This paper briefly describes the Cassini mission and spacecraft and, in somewhat more detail, the scientific instruments.

Performance Assessment of Two GPS Receivers on Space Shuttle

NASA Technical Reports Server (NTRS)

Schroeder, Christine A.; Schutz, Bob E.

1996-01-01

Space Shuttle STS-69 was launched on September 7, 1995, carrying the Wake Shield Facility (WSF-02) among its payloads. The mission included two GPS receivers: a Collins 3M receiver onboard the Endeavour and an Osborne flight TurboRogue, known as the TurboStar, onboard the WSF-02. Two of the WSF-02 GPS Experiment objectives were to: (1) assess the ability to use GPS in a relative satellite positioning mode using the receivers on Endeavour and WSF-02; and (2) assess the performance of the receivers to support high precision orbit determination at the 400 km altitude. Three ground tests of the receivers were conducted in order to characterize the respective receivers. The analysis of the tests utilized the Double Differencing technique. A similar test in orbit was conducted during STS-69 while the WSF-02 was held by the Endeavour robot arm for a one hour period. In these tests, biases were observed in the double difference pseudorange measurements, implying that biases up to 140 m exist which do not cancel in double differencing. These biases appear to exist in the Collins receiver, but their effect can be mitigated by including measurement bias parameters to accommodate them in an estimation process. An additional test was conducted in which the orbit of the combined Endeavour/WSF-02 was determined independently with each receiver. These one hour arcs were based on forming double differences with 13 TurboRogue receivers in the global IGS network and estimating pseudorange biases for the Collins. Various analyses suggest the TurboStar overall orbit accuracy is about one to two meters for this period, based on double differenced phase residuals of 34 cm. These residuals indicate the level of unmodeled forces on Endeavour produced by gravitational and nongravitational effects. The rms differences between the two independently determined orbits are better than 10 meters, thereby demonstrating the accuracy of the Collins-determined orbit at this level as well as the accuracy of the relative positioning using these two receivers.

The design of two-stage-to-orbit vehicles

NASA Technical Reports Server (NTRS)

1991-01-01

Two separate student design groups developed conceptual designs for a two-stage-to-orbit vehicle, with each design group consisting of a carrier team and an orbiter team. A two-stage-to-orbit system is considered in the event that single-stage-to-orbit is deemed not feasible in the foreseeable future; the two-stage system would also be used as a complement to an already existing heavy lift vehicle. The design specifications given are to lift a 10,000-lb payload 27 ft long by 10 ft diameter, to low Earth orbit (300 n.m.) using an air breathing carrier configuration that will take off horizontally within 15,000 ft. The staging Mach number and altitude were to be determined by the design groups. One group designed a delta wing/body carrier with the orbiter nested within the fuselage of the carrier, and the other group produced a blended cranked-delta wing/body carrier with the orbiter in the more conventional piggyback configuration. Each carrier used liquid hydrogen-fueled turbofanramjet engines, with data provided by General Electric Aircraft Engine Group. While one orbiter used a full-scale Space Shuttle Main Engine (SSME), the other orbiter employed a half-scale SSME coupled with scramjet engines, with data again provided by General Electric. The two groups conceptual designs, along with the technical trade-offs, difficulties, and details that surfaced during the design process are presented.

Mission Analysis Program for Solar Electric Propulsion (MAPSEP). Volume 1: Analytical manual for earth orbital MAPSEP

NASA Technical Reports Server (NTRS)

1975-01-01

An introduction to the MAPSEP organization and a detailed analytical description of all models and algorithms are given. These include trajectory and error covariance propagation methods, orbit determination processes, thrust modeling, and trajectory correction (guidance) schemes. Earth orbital MAPSEP contains the capability of analyzing almost any currently projected low thrust mission from low earth orbit to super synchronous altitudes. Furthermore, MAPSEP is sufficiently flexible to incorporate extended dynamic models, alternate mission strategies, and almost any other system requirement imposed by the user. As in the interplanetary version, earth orbital MAPSEP represents a trade-off between precision modeling and computational speed consistent with defining necessary system requirements. It can be used in feasibility studies as well as in flight operational support. Pertinent operational constraints are available both implicitly and explicitly. However, the reader should be warned that because of program complexity, MAPSEP is only as good as the user and will quickly succumb to faulty user inputs.

Proceedings of the 20th International Symposium on Space Flight Dynamics

NASA Technical Reports Server (NTRS)

Woodard, Mark (Editor); Stengle, Tom (Editor)

2007-01-01

Topics include: Measuring Image Navigation and Registration Performance at the 3-Sigma Level Using Platinum Quality Landmarks; Flight Dynamics Performances of the MetOp A Satellite during the First Months of Operations; Visual Navigation - SARE Mission; Determining a Method of Enabling and Disabling the Integral Torque in the SDO Science and Inertial Mode Controllers; Guaranteeing Pointing Performance of the SDO Sun-Pointing Controllers in Light of Nonlinear Effects; SDO Delta H Mode Design and Analysis; Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter; Broken-Plane Maneuver Applications for Earth to Mars Trajectories; ExoMars Mission Analysis and Design - Launch, Cruise and Arrival Analyses; Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance; Mars Reconnaissance Orbiter Interplanetary Cruise Navigation; Motion Parameters Determination of the SC and Phobos in the Project Phobos-Grunt; GRAS NRT Precise Orbit Determination: Operational Experience; Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods; Orbit Determination System for Low Earth Orbit Satellites; Precise Orbit Determination for ALOS; Anti-Collision Function Design and Performances of the CNES Formation Flying Experiment on the PRISMA Mission; CNES Approaching Guidance Experiment within FFIORD; Maneuver Recovery Analysis for the Magnetospheric Multiscale Mission; SIMBOL-X: A Formation Flying Mission on HEO for Exploring the Universe; Spaceborne Autonomous and Ground Based Relative Orbit Control for the TerraSAR-X/TanDEM-X Formation; First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations; Automated Target Planning for FUSE Using the SOVA Algorithm; Space Technology 5 Post-Launch Ground Attitude Estimation Experience; Standardizing Navigation Data: A Status Update; and A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-01-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

NASA Astrophysics Data System (ADS)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-10-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Orbit determination for ISRO satellite missions

NASA Astrophysics Data System (ADS)

Rao, Ch. Sreehari; Sinha, S. K.

Indian Space Research Organisation (ISRO) has been successful in using the in-house developed orbit determination and prediction software for satellite missions of Bhaskara, Rohini and APPLE. Considering the requirements of satellite missions, software packages are developed, tested and their accuracies are assessed. Orbit determination packages developed are SOIP, for low earth orbits of Bhaskara and Rohini missions, ORIGIN and ODPM, for orbits related to all phases of geo-stationary missions and SEGNIP, for drift and geo-stationary orbits. Software is tested and qualified using tracking data of SIGNE-3, D5-B, OTS, SYMPHONIE satellites with the help of software available with CNES, ESA and DFVLR. The results match well with those available from these agencies. These packages have supported orbit determination successfully throughout the mission life for all ISRO satellite missions. Member-Secretary

Quantum Storage of Three-Dimensional Orbital-Angular-Momentum Entanglement in a Crystal.

PubMed

Zhou, Zong-Quan; Hua, Yi-Lin; Liu, Xiao; Chen, Geng; Xu, Jin-Shi; Han, Yong-Jian; Li, Chuan-Feng; Guo, Guang-Can

2015-08-14

Here we present the quantum storage of three-dimensional orbital-angular-momentum photonic entanglement in a rare-earth-ion-doped crystal. The properties of the entanglement and the storage process are confirmed by the violation of the Bell-type inequality generalized to three dimensions after storage (S=2.152±0.033). The fidelity of the memory process is 0.993±0.002, as determined through complete quantum process tomography in three dimensions. An assessment of the visibility of the stored weak coherent pulses in higher-dimensional spaces demonstrates that the memory is highly reliable for 51 spatial modes. These results pave the way towards the construction of high-dimensional and multiplexed quantum repeaters based on solid-state devices. The multimode capacity of rare-earth-based optical processors goes beyond the temporal and the spectral degree of freedom, which might provide a useful tool for photonic information processing.

Comprehensive analysis of airborne contaminants from recent Spacelab missions

NASA Technical Reports Server (NTRS)

Matney, M. L.; Boyd, J. F.; Covington, P. A.; Leano, H. J.; Pierson, D. L.; Limero, T. F.; James, J. T.

1993-01-01

The Shuttle experiences unique air contamination problems because of microgravity and the closed environment. Contaminant build-up in the closed atmosphere and the lack of a gravitational settling mechanism have produced some concern in previous missions about the amount of solid and volatile airborne contaminants in the Orbiter and Spacelab. Degradation of air quality in the Orbiter/Spacelab environment, through processes such as chemical contamination, high solid-particulate levels, and high microbial levels, may affect crew performance and health. A comprehensive assessment of the Shuttle air quality was undertaken during STS-40 and STS-42 missions, in which a variety of air sampling and monitoring techniques were employed to determine the contaminant load by characterizing and quantitating airborne contaminants. Data were collected on the airborne concentrations of volatile organic compounds, microorganisms, and particulate matter collected on Orbiter/Spacelab air filters. The results showed that STS-40/42 Orbiter/Spacelab air was toxicologically safe to breathe, except during STS-40 when the Orbiter Refrigerator/Freezer unit was releasing noxious gases in the middeck. On STS-40, the levels of airborne bacteria appeared to increase as the mission progressed; however, this trend was not observed for the STS-42 mission. Particulate matter in the Orbiter/Spacelab air filters was chemically analyzed in order to determine the source of particles. Only small amounts of rat hair and food bar (STS-40) and traces of soiless medium (STS-42) were detected in the Spacelab air filters, indicating that containment for Spacelab experiments was effective.

General view of the Orbiter Discovery on runway 33 at ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the Orbiter Discovery on runway 33 at Kennedy Space Center shortly after landing. The orbiter is processed and prepared for being towed to the Orbiter Processing Facility for continued post flight processing and pre flight preparations for its next mission. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Independent Orbiter Assessment (IOA): Analysis of the active thermal control subsystem

NASA Technical Reports Server (NTRS)

Sinclair, S. K.; Parkman, W. E.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical (PCIs) items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results corresponding to the Orbiter Active Thermal Control Subsystem (ATCS) are documented. The major purpose of the ATCS is to remove the heat, generated during normal Shuttle operations from the Orbiter systems and subsystems. The four major components of the ATCS contributing to the heat removal are: Freon Coolant Loops; Radiator and Flow Control Assembly; Flash Evaporator System; and Ammonia Boiler System. In order to perform the analysis, the IOA process utilized available ATCS hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 310 failure modes analyzed, 101 were determined to be PCIs.

A stellar tracking reference system

NASA Technical Reports Server (NTRS)

Klestadt, B.

1971-01-01

A stellar attitude reference system concept for satellites was studied which promises to permit continuous precision pointing of payloads with accuracies of 0.001 degree without the use of gyroscopes. It is accomplished with the use of a single, clustered star tracker assembly mounted on a non-orthogonal, two gimbal mechanism, driven so as to unwind satellite orbital and orbit precession rates. A set of eight stars was found which assures the presence of an adequate inertial reference on a continuous basis in an arbitrary orbit. Acquisition and operational considerations were investigated and inherent reference redundancy/reliability was established. Preliminary designs for the gimbal mechanism, its servo drive, and the star tracker cluster with its associated signal processing were developed for a baseline sun-synchronous, noon-midnight orbit. The functions required of the onboard computer were determined and the equations to be solved were found. In addition detailed error analyses were carried out, based on structural, thermal and other operational considerations.

Optimization of Insertion Cost for Transfer Trajectories to Libration Point Orbits

NASA Technical Reports Server (NTRS)

Howell, K. C.; Wilson, R. S.; Lo, M. W.

1999-01-01

The objective of this work is the development of efficient techniques to optimize the cost associated with transfer trajectories to libration point orbits in the Sun-Earth-Moon four body problem, that may include lunar gravity assists. Initially, dynamical systems theory is used to determine invariant manifolds associated with the desired libration point orbit. These manifolds are employed to produce an initial approximation to the transfer trajectory. Specific trajectory requirements such as, transfer injection constraints, inclusion of phasing loops, and targeting of a specified state on the manifold are then incorporated into the design of the transfer trajectory. A two level differential corrections process is used to produce a fully continuous trajectory that satisfies the design constraints, and includes appropriate lunar and solar gravitational models. Based on this methodology, and using the manifold structure from dynamical systems theory, a technique is presented to optimize the cost associated with insertion onto a specified libration point orbit.

Space Trajectory Error Analysis Program (STEAP) for halo orbit missions. Volume 2: Programmer's manual

NASA Technical Reports Server (NTRS)

Byrnes, D. V.; Carney, P. C.; Underwood, J. W.; Vogt, E. D.

1974-01-01

The six month effort was responsible for the development, test, conversion, and documentation of computer software for the mission analysis of missions to halo orbits about libration points in the earth-sun system. The software consisting of two programs called NOMNAL and ERRAN is part of the Space Trajectories Error Analysis Programs. The program NOMNAL targets a transfer trajectory from earth on a given launch date to a specified halo orbit on a required arrival date. Either impulsive or finite thrust insertion maneuvers into halo orbit are permitted by the program. The transfer trajectory is consistent with a realistic launch profile input by the user. The second program ERRAN conducts error analyses of the targeted transfer trajectory. Measurements including range, doppler, star-planet angles, and apparent planet diameter are processed in a Kalman-Schmidt filter to determine the trajectory knowledge uncertainty.

Mapping Vesta: First Results from Dawn's Survey Orbit

NASA Technical Reports Server (NTRS)

Jaumann, R.; Yingst, A. R.; Pieters, C. M.; Russell, C. T.; Raymond, C. A.; Neukum, G.; Mottola, S.; Keller, H. U.; Nathues, A.; Sierks, H.;

Drag Coefficient Estimation in Orbit Determination

NASA Astrophysics Data System (ADS)

McLaughlin, Craig A.; Manee, Steve; Lichtenberg, Travis

2011-07-01

Drag modeling is the greatest uncertainty in the dynamics of low Earth satellite orbits where ballistic coefficient and density errors dominate drag errors. This paper examines fitted drag coefficients found as part of a precision orbit determination process for Stella, Starlette, and the GEOSAT Follow-On satellites from 2000 to 2005. The drag coefficients for the spherical Stella and Starlette satellites are assumed to be highly correlated with density model error. The results using MSIS-86, NRLMSISE-00, and NRLMSISE-00 with dynamic calibration of the atmosphere (DCA) density corrections are compared. The DCA corrections were formulated for altitudes of 200-600 km and are found to be inappropriate when applied at 800 km. The yearly mean fitted drag coefficients are calculated for each satellite for each year studied. The yearly mean drag coefficients are higher for Starlette than Stella, where Starlette is at a higher altitude. The yearly mean fitted drag coefficients for all three satellites decrease as solar activity decreases after solar maximum.

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.

Evaluation of orbits with incomplete knowledge of the mathematical expectancy and the matrix of covariation of errors

NASA Technical Reports Server (NTRS)

Bakhshiyan, B. T.; Nazirov, R. R.; Elyasberg, P. E.

1980-01-01

The problem of selecting the optimal algorithm of filtration and the optimal composition of the measurements is examined assuming that the precise values of the mathematical expectancy and the matrix of covariation of errors are unknown. It is demonstrated that the optimal algorithm of filtration may be utilized for making some parameters more precise (for example, the parameters of the gravitational fields) after preliminary determination of the elements of the orbit by a simpler method of processing (for example, the method of least squares).

Infrared Spectroscopy of Symbiotic Stars. II. Orbits for Five S-Type Systems with Two-Year Periods

NASA Astrophysics Data System (ADS)

Fekel, Francis C.; Hinkle, Kenneth H.; Joyce, Richard R.; Skrutskie, Michael F.

2000-12-01

Infrared radial velocities have been used to determine orbital elements for the cool giants of five well-known symbiotic systems, Z And, AG Dra, V443 Her, AX Per, and FG Ser, all of which have orbital periods near the two-year mean period for S-type symbiotics. The new orbits are in general agreement with previous orbits derived from optical velocities. From the combined optical and infrared velocities, improved orbital elements for the five systems have been determined. Each of the orbital periods has been determined solely from the radial-velocity data. The orbits are circular and have quite small mass functions of 0.001-0.03 Msolar. The infrared velocities of AG Dra do not show the large orbital velocity residuals found for its optical radial velocities.

Independent Orbiter Assessment (IOA): Analysis of the pyrotechnics subsystem

NASA Technical Reports Server (NTRS)

Robinson, W. W.

1988-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Pyrotechnics hardware. The IOA analysis process utilized available pyrotechnics hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

KSC-04PD-1813

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. In the Orbiter Processing Facility, an External Tank (ET) digital still camera is positioned into the right-hand liquid oxygen umbilical well on Space Shuttle Atlantis to determine if it fits properly. NASA is pursuing use of the camera, beginning with the Shuttles Return To Flight, to obtain and downlink high-resolution images of the ET following separation of the ET from the orbiter after launch. The Kodak camera will record 24 images, at one frame per 1.5 seconds, on a flash memory card. After orbital insertion, the crew will transfer the images from the memory card to a laptop computer. The files will then be downloaded through the Ku-band system to the Mission Control Center in Houston for analysis.

KSC-04pd1813

NASA Image and Video Library

2004-09-17

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an External Tank (ET) digital still camera is positioned into the right-hand liquid oxygen umbilical well on Space Shuttle Atlantis to determine if it fits properly. NASA is pursuing use of the camera, beginning with the Shuttle’s Return To Flight, to obtain and downlink high-resolution images of the ET following separation of the ET from the orbiter after launch. The Kodak camera will record 24 images, at one frame per 1.5 seconds, on a flash memory card. After orbital insertion, the crew will transfer the images from the memory card to a laptop computer. The files will then be downloaded through the Ku-band system to the Mission Control Center in Houston for analysis.

Signal design study for shuttle/TDRSS Ku-band uplink

NASA Technical Reports Server (NTRS)

1976-01-01

The adequacy of the signal design approach chosen for the TDRSS/orbiter uplink was evaluated. Critical functions and/or components associated with the baseline design were identified, and design alternatives were developed for those areas considered high risk. A detailed set of RF and signal processing performance specifications for the orbiter hardware associated with the TDRSS/orbiter Ku band uplink was analyzed. Performances of a detailed design of the PN despreader, the PSK carrier synchronization loop, and the symbol synchronizer are identified. The performance of the downlink signal by means of computer simulation to obtain a realistic determination of bit error rate degradations was studied. The three channel PM downlink signal was detailed by means of analysis and computer simulation.

Lunar Prospector Orbit Determination Uncertainties Using the High Resolution Lunar Gravity Models

NASA Technical Reports Server (NTRS)

Carranza, Eric; Konopliv, Alex; Ryne, Mark

1999-01-01

The Lunar Prospector (LP) mission began on January 6, 1998, when the LP spacecraft was launched from Cape Canaveral, Florida. The objectives of the mission were to determine whether water ice exists at the lunar poles, generate a global compositional map of the lunar surface, detect lunar outgassing, and improve knowledge of the lunar magnetic and gravity fields. Orbit determination of LP performed at the Jet Propulsion Laboratory (JPL) is conducted as part of the principal science investigation of the lunar gravity field. This paper will describe the JPL effort in support of the LP Gravity Investigation. This support includes high precision orbit determination, gravity model validation, and data editing. A description of the mission and its trajectory will be provided first, followed by a discussion of the orbit determination estimation procedure and models. Accuracies will be examined in terms of orbit-to-orbit solution differences, as a function of oblateness model truncation, and inclination in the plane-of-sky. Long term predictions for several gravity fields will be compared to the reconstructed orbits to demonstrate the accuracy of the orbit determination and oblateness fields developed by the Principal Gravity Investigator.

Independent Orbiter Assessment (IOA): Analysis of the DPS subsystem

NASA Technical Reports Server (NTRS)

Lowery, H. J.; Haufler, W. A.; Pietz, K. C.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis/Critical Items List (FMEA/CIL) is presented. The IOA approach features a top-down analysis of the hardware to independently determine failure modes, criticality, and potential critical items. The independent analysis results corresponding to the Orbiter Data Processing System (DPS) hardware are documented. The DPS hardware is required for performing critical functions of data acquisition, data manipulation, data display, and data transfer throughout the Orbiter. Specifically, the DPS hardware consists of the following components: Multiplexer/Demultiplexer (MDM); General Purpose Computer (GPC); Multifunction CRT Display System (MCDS); Data Buses and Data Bus Couplers (DBC); Data Bus Isolation Amplifiers (DBIA); Mass Memory Unit (MMU); and Engine Interface Unit (EIU). The IOA analysis process utilized available DPS hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Due to the extensive redundancy built into the DPS the number of critical items are few. Those identified resulted from premature operation and erroneous output of the GPCs.

First independent lunar gravity field solution in the framework of project GRAZIL

NASA Astrophysics Data System (ADS)

Wirnsberger, Harald; Krauss, Sandro; Klinger, Beate; Mayer-Gürr, Torsten

2017-04-01

The twin satellite mission Gravity Recovery and Interior Laboratory (GRAIL) aims to recovering the lunar gravity field by means of intersatellite Ka-band ranging (KBR) observations. In order to exploit the potential of KBR data, absolute position information of the two probes is required. Hitherto, the Graz lunar gravity field models (GrazLGM) relies on the official orbit products provided by NASA. In this contribution, we present for the first time a completely independent Graz lunar gravity field model to spherical harmonic degree and order 420. The reduced dynamic orbits of the two probes are determined using variational equations following a batch least squares differential adjustment process. These orbits are based on S-band radiometric tracking data collected by the Deep Space Network and are used for the independent GRAIL gravity field recovery. To reveal a highly accurate lunar gravity field, an integral equation approach using short orbital arcs is adopted to process the KBR data. A comparison to state-of-the-art lunar gravity models computed at NASA-GSFC, NASA-JPL and AIUB demonstrate the progress of Graz lunar gravity field models derived within the project GRAZIL.

Spin Evolution of Stellar Progenitors in Compact Binaries

NASA Astrophysics Data System (ADS)

Steinle, Nathan; Kesden, Michael

2018-01-01

Understanding the effects of various processes on the spins of stellar progenitors in compact binary systems is important for modeling the binary’s evolution and thus for interpreting the gravitational radiation emitted during inspiral and merger. Tides, winds, and natal kicks can drastically modify the binary parameters: tidal interactions increase the spin magnitudes, align the spins with the orbital angular momentum, and circularize the orbit; stellar winds decrease the spin magnitudes and cause mass loss; and natal kicks can misalign the spins and orbital angular momentum or even disrupt the binary. Also, during Roche lobe overflow, the binary may experience either stable mass transfer or common envelope evolution. The former can lead to a mass ratio reversal and alter the component spins, while the latter can dramatically shrink the binary separation. For a wide range of physically reasonable stellar-evolution scenarios, we compare the timescales of these processes to assess their relative contributions in determining the initial spins of compact binary systems.

Evaluation of TDRSS-user orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Hodjatzadeh, M.; Samii, M. V.; Doll, C. E.; Hart, R. C.; Mistretta, G. D.

1991-01-01

The development of the Real-Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination on a Disk Operating System (DOS) based Personal Computer (PC) is addressed. The results of a study to compare the orbit determination accuracy of a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOD/E with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), is addressed. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for the Earth Radiation Budget Satellite (ERBS); the maximum solution differences were less than 25 m after the filter had reached steady state.

Space shuttle orbiter digital data processing system timing sensitivity analysis OFT ascent phase

NASA Technical Reports Server (NTRS)

Lagas, J. J.; Peterka, J. J.; Becker, D. A.

1977-01-01

Dynamic loads were investigated to provide simulation and analysis of the space shuttle orbiter digital data processing system (DDPS). Segments of the ascent test (OFT) configuration were modeled utilizing the information management system interpretive model (IMSIM) in a computerized simulation modeling of the OFT hardware and software workload. System requirements for simulation of the OFT configuration were defined, and sensitivity analyses determined areas of potential data flow problems in DDPS operation. Based on the defined system requirements and these sensitivity analyses, a test design was developed for adapting, parameterizing, and executing IMSIM, using varying load and stress conditions for model execution. Analyses of the computer simulation runs are documented, including results, conclusions, and recommendations for DDPS improvements.

Late stages of accumulation and early evolution of the planets

NASA Technical Reports Server (NTRS)

Vityazev, Andrey V.; Perchernikova, G. V.

1991-01-01

Recently developed solutions of problems are discussed that were traditionally considered fundamental in classical solar system cosmogony: determination of planetary orbit distribution patterns, values for mean eccentricity and orbital inclinations of the planets, and rotation periods and rotation axis inclinations of the planets. Two important cosmochemical aspects of accumulation are examined: the time scale for gas loss from the terrestrial planet zone, and the composition of the planets in terms of isotope data. It was concluded that the early beginning of planet differentiation is a function of the heating of protoplanets during collisions with large (thousands of kilometers) bodies. Energetics, heat mass transfer processes, and characteristic time scales of these processes at the early stages of planet evolution are considered.

Field Geology/Processes

NASA Technical Reports Server (NTRS)

Allen, Carlton; Jakes, Petr; Jaumann, Ralf; Marshall, John; Moses, Stewart; Ryder, Graham; Saunders, Stephen; Singer, Robert

1996-01-01

The field geology/process group examined the basic operations of a terrestrial field geologist and the manner in which these operations could be transferred to a planetary lander. Four basic requirements for robotic field geology were determined: geologic content; surface vision; mobility; and manipulation. Geologic content requires a combination of orbital and descent imaging. Surface vision requirements include range, resolution, stereo, and multispectral imaging. The minimum mobility for useful field geology depends on the scale of orbital imagery. Manipulation requirements include exposing unweathered surfaces, screening samples, and bringing samples in contact with analytical instruments. To support these requirements, several advanced capabilities for future development are recommended. Capabilities include near-infrared reflectance spectroscopy, hyper-spectral imaging, multispectral microscopy, artificial intelligence in support of imaging, x ray diffraction, x ray fluorescence, and rock chipping.

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.

A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver.

PubMed

Wang, Fuhong; Gong, Xuewen; Sang, Jizhang; Zhang, Xiaohong

2015-12-04

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.

Physics Notes.

ERIC Educational Resources Information Center

School Science Review, 1982

1982-01-01

Discusses determination of elliptical path of a satellite caught into orbit by the sun or earth; using microcomputer as signal generator (includes program listing); collision process; simple hysteresis loop using double beam CRO; method of demonstrating parallelogram of forces; measuring radius of electron beam curvature; and half-life of thorium…

Autonomous Navigation for Deep Space Missions

NASA Technical Reports Server (NTRS)

Bhaskaran, Shyam

2012-01-01

Navigation (determining where the spacecraft is at any given time, controlling its path to achieve desired targets), performed using ground-in- the-loop techniques: (1) Data includes 2-way radiometric (Doppler, range), interferometric (Delta- Differential One-way Range), and optical (images of natural bodies taken by onboard camera) (2) Data received on the ground, processed to determine orbit, commands sent to execute maneuvers to control orbit. A self-contained, onboard, autonomous navigation system can: (1) Eliminate delays due to round-trip light time (2) Eliminate the human factors in ground-based processing (3) Reduce turnaround time from navigation update to minutes, down to seconds (4) React to late-breaking data. At JPL, we have developed the framework and computational elements of an autonomous navigation system, called AutoNav. It was originally developed as one of the technologies for the Deep Space 1 mission, launched in 1998; subsequently used on three other spacecraft, for four different missions. The primary use has been on comet missions to track comets during flybys, and impact one comet.

Mars approach navigation using Doppler and range measurements to surface beacons and orbiting spacecraft

NASA Technical Reports Server (NTRS)

Thurman, Sam W.; Estefan, Jeffrey A.

1991-01-01

Approximate analytical models are developed and used to construct an error covariance analysis for investigating the range of orbit determination accuracies which might be achieved for typical Mars approach trajectories. The sensitivity or orbit determination accuracy to beacon/orbiter position errors and to small spacecraft force modeling errors is also investigated. The results indicate that the orbit determination performance obtained from both Doppler and range data is a strong function of the inclination of the approach trajectory to the Martian equator, for surface beacons, and for orbiters, the inclination relative to the orbital plane. Large variations in performance were also observed for different approach velocity magnitudes; Doppler data in particular were found to perform poorly in determining the downtrack (along the direction of flight) component of spacecraft position. In addition, it was found that small spacecraft acceleration modeling errors can induce large errors in the Doppler-derived downtrack position estimate.

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (left) listens to Kathy Laufenberg, Orbiter Airframe Engineering ground area manager, with United Space Alliance, about corrosion work being done on the external tank door of orbiter Endeavour. On either side of Laufenberg are Tom Roberts, Airframe Engineering System specialist, also with USA, and Joy Huff, with KSC Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

NASA Image and Video Library

2004-02-25

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (left) listens to Kathy Laufenberg, Orbiter Airframe Engineering ground area manager, with United Space Alliance, about corrosion work being done on the external tank door of orbiter Endeavour. On either side of Laufenberg are Tom Roberts, Airframe Engineering System specialist, also with USA, and Joy Huff, with KSC Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

An Analysis of the Space Transportation System Launch Rate Capability Utilizing Q-GERT Simulation Techniques.

DTIC Science & Technology

1982-12-01

VAPE was modeled to determine this launch rate and to determine the processing times for an Orbiter at VAPe . This informa- 21 tion was then used in the...year (node 79 and activity ?1). ETa are then selected to be sent to either KSC or VAPE (node 80). This decision is made (using Ur 8) on the basis of

Satellite orbit determination using quantum correlation technology

NASA Astrophysics Data System (ADS)

Zhang, Bo; Sun, Fuping; Zhu, Xinhui; Jia, Xiaolin

2018-03-01

After the presentation of second-order correlation ranging principles with quantum entanglement, the concept of quantum measurement is introduced to dynamic satellite precise orbit determination. Based on the application of traditional orbit determination models for correcting the systematic errors within the satellite, corresponding models for quantum orbit determination (QOD) are established. This paper experiments on QOD with the BeiDou Navigation Satellite System (BDS) by first simulating quantum observations of 1 day arc-length. Then the satellite orbits are resolved and compared with the reference precise ephemerides. Subsequently, some related factors influencing the accuracy of QOD are discussed. Furthermore, the accuracy for GEO, IGSO and MEO satellites increase about 20, 30 and 10 times, respectively, compared with the results from the resolution by measured data. Therefore, it can be expected that quantum technology may also bring delightful surprises to satellite orbit determination as have already emerged in other fields.

Computer processing of Mars Odyssey THEMIS IR imaging, MGS MOLA altimetry and Mars Express stereo imaging to locate Airy-0, the Mars prime meridian reference

NASA Astrophysics Data System (ADS)

Duxbury, Thomas; Neukum, Gerhard; Smith, David E.; Christensen, Philip; Neumann, Gregory; Albee, Arden; Caplinger, Michael; Seregina, N. V.; Kirk, Randolph L.

The small crater Airy-0 was selected from Mariner 9 images to be the reference for the Mars prime meridian. Initial analyses were made in year 2000 to tie Viking Orbiter and Mars Orbiter Camera images of Airy-0 to the evolving Mars Orbiter Laser Altimeter global digital terrain model to improve the location accuracy of Airy-0. Based upon this tie and radiometric tracking of landers / rovers from earth, new expressions for the Mars spin axis direction, spin rate and prime meridian epoch value were produced to define the orientation of the Martian surface in inertial space over time. Now that the Mars Global Surveyor mission and the Mars Orbiter Laser Altimeter global digital terrain model are complete, a more exhaustive study has been performed to determine the location of Airy-0 relative to the global terrain grid. THEMIS IR image cubes of the Airy and Gale crater regions were tied to the global terrain grid using precision stereo photogrammetric image processing techniques. The Airy-0 location was determined to be within 50 meters of the currently defined IAU prime meridian, with this offset at the limiting absolute accuracy of the global terrain grid. Additional outputs of this study were a controlled multi-band photomosaic of Airy, precision alignment and geometric models of the ten THEMIS IR bands and a controlled multi-band photomosaic of Gale crater used to validate the Mars Surface Laboratory operational map products supporting their successful landing on Mars.

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.

On-Orbit Ephemeris Determination with Radio Doppler Validation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Dallmann, Nicholas; Proicou, Michael Chris; Seitz, Daniel Nathan

2016-02-09

Multiple CubeSats are often released from the same host spacecraft into virtually the same orbit at nearly the same time. A satellite team needs the ability to identify and track its own satellites as soon as possible. However, this can be a difficult and confusing task with a large number of satellites. Los Alamos National Laboratory encountered this issue during a launch of LANL-designed CubeSats that were released with more than 20 other objects. A simple radio Doppler method used shortly after launch by the Los Alamos team to select its satellites of interest from the list of available trackedmore » ephemerides is described. This method can also be used for automated real time ephemeris validation. For future efforts, each LANL-designed CubeSat will automatically perform orbit determination from the position, velocity, and covariance estimates provided by an added on-board GPS receiver. This self-determined ephemeris will be automatically downlinked by ground stations for mission planning, antenna tracking, Doppler-pre-correction, etc. A simple algorithm based on established theory and well suited for embedded on-board processing is presented. The trades examined in selecting the algorithm components and data formats are briefly discussed, as is the expected performance.« less

Fabrication process scale-up and optimization for a boron-aluminum composite radiator

NASA Technical Reports Server (NTRS)

Okelly, K. P.

1973-01-01

Design approaches to a practical utilization of a boron-aluminum radiator for the space shuttle orbiter are presented. The program includes studies of laboratory composite material processes to determine the feasibility of a structural and functional composite radiator panel, and to estimate the cost of its fabrication. The objective is the incorporation of boron-aluminum modulator radiator on the space shuttle.

ISTP SBIR phase 1 Full-Sky Scanner: A feasibility study

NASA Technical Reports Server (NTRS)

1986-01-01

The objective was to develop a Full-Sky Sensor (FSS) to detect the Earth, Sun and Moon from a spinning spacecraft. The concept adopted has infinitely variable resolution. A high-speed search mode is implemented on the spacecraft. The advantages are: (1) a single sensor determines attitude parameters from Earth, Sun and Moon, thus eliminating instrument mounting errors; (2) the bias between the actual spacecraft spin axis and the intended spin axis can be determined; (3) cost is minimized; and (4) ground processing is straightforward. The FSS is a modification of an existing flight-proven sensor. Modifications to the electronics are necessary to accommodate the amplitude range and signal width range of the celestial bodies to be detected. Potential applications include ISTP missions, Multi-Spacecraft Satellite Program (MSSP), dual-spin spacecraft at any altitude, spinning spacecraft at any altitude, and orbit parameter determination for low-Earth orbits.

ISTP SBIR phase 1 Full-Sky Scanner: A feasibility study

NASA Astrophysics Data System (ADS)

1986-08-01

The objective was to develop a Full-Sky Sensor (FSS) to detect the Earth, Sun and Moon from a spinning spacecraft. The concept adopted has infinitely variable resolution. A high-speed search mode is implemented on the spacecraft. The advantages are: (1) a single sensor determines attitude parameters from Earth, Sun and Moon, thus eliminating instrument mounting errors; (2) the bias between the actual spacecraft spin axis and the intended spin axis can be determined; (3) cost is minimized; and (4) ground processing is straightforward. The FSS is a modification of an existing flight-proven sensor. Modifications to the electronics are necessary to accommodate the amplitude range and signal width range of the celestial bodies to be detected. Potential applications include ISTP missions, Multi-Spacecraft Satellite Program (MSSP), dual-spin spacecraft at any altitude, spinning spacecraft at any altitude, and orbit parameter determination for low-Earth orbits.

Pathogenesis of ophthalmopathy in autoimmune thyroid disease.

PubMed

Heufelder, A E

2000-01-01

What causes GO is still a mystery, but the disease process results from a complex interplay of genetic and environmental factors. Genes such as those for HLA genes may determine a patient's susceptibility to the disease and its severity, but environmental factors, often unknown, may determine its course. Once established, the chronic inflammatory process within the orbital tissues appears to take on a momentum of its own. Given our current state of knowledge, the following working scheme for the pathogenesis of GO can be proposed (Fig. 1): On the background of a permissive immunogenetic milieu, circulating T cells in patients with GD, directed against certain antigens on thyroid follicular cells, recognize antigenic epitopes that are shared by tissues contained in the orbital space. Of the cell types residing in these tissues, preadipocytes and fibroblasts, most likely act as target and effector cells of the orbital immune process, respectively. This includes preadipocyte fibroblasts present in the perimysium of extraocular muscles, which do not appear to be immunologically different from those located in the orbital connective tissue. Orbital preadipocyte fibroblasts may be stimulated by unknown circulating or locally produced factors to differentiate into mature adipocytes that express increased levels of TSHr. How autoreactive T cells escape deletion and control by the immune system and come to be directed against a self-antigen presented by cells residing in the thyroid gland and extrathyroidal locations, is still unknown. Proliferation and expansion of autoreactive T cell clones may be due to mimicry of a host antigen by a microorganism, but this remains speculative. T cell recruitment into the orbital tissues is facilitated by certain chemokines and cytokines, which help to attract T cells by stimulating the expression of certain adhesion molecules (e.g., ICAM-1, VCAM-1, CD44) in vascular endothelium and connective tissue cells. These adhesion receptors are known to also play an important costimulatory role by activating T cells and facilitating antigen recognition, which amplifies the cellular immune process. Analysis of variable region genes of T cell antigen receptors in orbital T cells of patients with active GO has revealed their restricted TcR V gene usage, suggesting that antigen-driven selection and/or expansion of specific T cells may occur early in the evolution of GO. T cells and macrophages populating the orbital space are known to synthesize and release a [figure: see text] number of cytokines (most likely a Th1-type spectrum) into the surrounding tissue. Cytokines, oxygen free radicals and fibrogenic growth factors, released both from infiltrating inflammatory and residential cells, act upon orbital preadipocytes in a paracrine and autocrine manner to stimulate adipogenesis, fibroblast proliferation, glycosaminoglycan synthesis, and the expression of immunomodulatory molecules. Smoking, a well-known aggravating factor in GO, may aggravate tissue hypoxia and exert important immunomodulatory effects. The long held hypothesis of a thyroid cross-reactive antigen within the orbital tissues has recently gained significant support by an animal model of GO, and by in vitro and ex vivo studies. If confirmed in immunological studies, these data may well explain the localized infiltration of the orbital tissues by autoreactive lymphocytes that share intriguing molecular features with intrathyroidal lymphocytes. Local release of particular cytokines, TSHr-directed antibodies, or other factors might further enhance adipogenesis, glycosaminoglycan synthesis and expression of immunomodulatory proteins within the orbit. Other factors, including inflammatory cytokines, might act as counterbalancing inhibitors of these effects. However, if the net effect of these changes is to increase the volume of the fatty connective tissues within the orbit, then proptosis, extraocular muscle dysfunction, and periorbital congestion will ensue. Whether this hypothetical sequence of events will finally explain the involvement of the orbit in GD is unknown. Future studies will be aimed at identifying factors that might modulate adipogenesis in orbital cells and clarifying the link between adipogenesis and TSHr expression in the orbit. Taken together, a number of important details in the complex pathogenesis of GO have been resolved in recent years, but many challenges are still ahead. Elucidation of the primary antigen and how it is recognized by the immune system will be key issues.

General view from inside the payload bay of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view from inside the payload bay of the Orbiter Discovery approximately along its centerline looking aft towards the bulkhead of the aft fuselage. Note panels and insulation removed for access to the orbiter's subsystems for inspection and post-mission processing. This photo was taken during the processing of the Orbiter Discovery after its final mission and in preparation for its transition to the National Air and Space Museum. This view was taken in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Evolution and Implementation of the NASA Robotic Conjunction Assessment Risk Analysis Concept of Operations

NASA Astrophysics Data System (ADS)

Newman, L.; Hejduk, M.; Frigm, R.; Duncan, M.

2014-09-01

On-orbit collisions pose a significant mission risk to satellites operating in the space environment. Recognizing the likelihood and consequence of on-orbit collisions, NASA has taken several proactive measures to mitigate the risk of both a catastrophic loss of mission and the increase in the space debris population. In fall 2004, NASA GSFC established an Agency-wide, institutionalized process and service for identifying and reacting to predicted close approaches. The team responsible for executing this mission is the NASA Robotic Conjunction Assessment Risk Analysis (CARA) team. By fall 2005, this process had resulted in the execution of the first collision avoidance maneuver by a NASA unmanned satellite. In February 2008, NASA adopted a policy, documented in NASA Procedural Requirement 8715.6a Process for Limiting Orbital Debris that directed maneuverable satellites to have such an on-orbit collision mitigation process. In 2009, NASA decided to require support for all operational satellites. By January 2014, the CARA team has processed nearly 500,000 close approach messages from the Joint Space Operations Center (JSpOC) and has assisted our mission customers with planning and executing over 75 collision avoidance maneuvers for unmanned satellites in LEO, GEO, and HEO orbital regimes. With the increase in number of operational missions supported; growth in the orbital debris environment due to events such as the intentional destruction of the Fengyun 1-C satellite in 2007 and collision between Iridium-33 and Cosmos-2251; and improvements to the United States Space Surveillance Network (SSN) and its ability to track, catalog, and screen against small debris objects, the demands on the CARA process have consequently required the CARA Concept of Operations (CONOPS) to evolve to manage those demands. This evolution is centered on the ability to effectively and efficiently manage JSpOC, CARA, and Mission Operations resources, applying operational and analytical efforts for conjunction events that pose significant collision risk and rapidly discarding conjunction events that do not. While the overall CARA methodology is largely unaffected, this CONOPS evolution manifests itself in several aspects of the CARA process: required data and information, communication of those data and information, and courses of actions based on those data and information. The changes affect all relevant stakeholders, including the CARA team at NASA GSFC, GSFC-dedicated Orbital Safety Analysts at the JSpOC, and Mission Operations flight teams and management. In each step of the CARA process, the CONOPS ensures that necessary (whether situational or actionable) information be sent to stakeholders to facilitate an effective and efficient management of resources and appropriate protection of data. The most significant paradigm shift is the movement to risk-based reporting. Since the consequence of the on-orbit collision scenario can be catastrophic, the CARA risk-based framework hinges on the collision probability, Pc, as the encapsulation of collision risk. This CONOPS characterizes collision risk as Red (high collision risk), Yellow (potential for becoming a high collision risk), or Green (low collision risk) based on the operationally-computed Pc. Using this risk characterization schema, the amount and content of conjunction information and analyses is determined and communicated to mission stakeholders. Major technical analyses that have been conducted in support of this CONOPS include defining risk-based thresholds for red, yellow, and green criteria; determining when conjunction-related information may not be mature enough to be actionable; and accounting for uncertainties in all the inputs to the process so that a nuanced assessment of risk can be made. This paper summarizes the analyses executed and decisions rendered during the implementation of this evolved CONOPS. Historical conjunction events of note are used as example scenarios of each risk characterization.

A Journey with MOM

NASA Technical Reports Server (NTRS)

Helfrich, Cliff; Berry, David S.; Bhat, Ramachandra; Border, James; Graat, Eric; Halsell, Allen; Kruizinga, Gerhard; Lau, Eunice; Mottinger, Neil; Rush, Brian;

TDRS orbit determination by radio interferometry

NASA Technical Reports Server (NTRS)

Pavloff, Michael S.

1994-01-01

In support of a NASA study on the application of radio interferometry to satellite orbit determination, MITRE developed a simulation tool for assessing interferometry tracking accuracy. The Orbit Determination Accuracy Estimator (ODAE) models the general batch maximum likelihood orbit determination algorithms of the Goddard Trajectory Determination System (GTDS) with the group and phase delay measurements from radio interferometry. ODAE models the statistical properties of tracking error sources, including inherent observable imprecision, atmospheric delays, clock offsets, station location uncertainty, and measurement biases, and through Monte Carlo simulation, ODAE calculates the statistical properties of errors in the predicted satellites state vector. This paper presents results from ODAE application to orbit determination of the Tracking and Data Relay Satellite (TDRS) by radio interferometry. Conclusions about optimal ground station locations for interferometric tracking of TDRS are presented, along with a discussion of operational advantages of radio interferometry.

[Interactions of fibroblasts, adipocytes and immunocompeent cells in the pathogenesis of endocrine ophthalmopathy].

PubMed

Heufelder, A E

2001-01-01

Graves' ophthalmopathy is thought to result from a complex interplay of genetic and environmental factors. Various genes including those coding for HLA may determine a patient's susceptibility to the disease and its severity, but in addition numerous and often unknown environmental factors may determine its course. The orbital immune process is thought to be initiated, on the background of a permissive immunogenetic milieu, by circulating T cells directed against certain antigens on thyroid follicular cells that also recognize antigenic epitopes which are shared by tissues contained in the orbital space. Analysis of variable region genes of T cell antigen receptors in orbital T cells of patients with active Graves' ophthalmopathy has revealed limited variability of TcR V gene usage, suggesting that antigen-driven selection and/or expansion of specific T cells may occur during the early stages of Graves' ophthalmopathy. T cell recruitment into the orbital tissues is facilitated by certain chemokines and cytokines, which attract T cells by stimulating the expression of several adhesion molecules (e.g. ICAM-1, VCAM-1, CD44) in vascular endothelium and connective tissue cells. Adhesion molecules are known to be important for a variety of interactions between immunocompetent cells, preadipocyte fibroblasts and adipocytes. In addition, these molecules play a central role in lymphocyte activation and localization, facilitating antigen recognition, T cell costimulation, and various effector-target cell functions at the inflammatory sites, which result in amplification of the cellular immune process in active Graves' ophthalmopathy. T cells and macrophages populate the orbital space and release a number of cytokines (most likely a Th-1-type spectrum) into the surrounding tissues. Cytokines, oxygen free radicals and fibrogenic growth factors, released both from infiltrating inflammatory and residential cells, act upon orbital preadipocytes in a paracrine and autocrine manner to stimulate adipogenesis, fibroblast proliferation, glycosaminoglycan synthesis, and the expression of immunomodulatory molecules. Smoking, a well-known aggravating factor in Graves' ophthalmopathy, may aggravate tissue hypoxia and exert important immunomodulatory and pro-oxidant effects. Differentiation of orbital preadipocyte fibroblasts into mature adipocytes expressing increased levels of TSHR may also be driven by stimulation with circulating or locally produced cytokines or effectors. TSHR-directed autoantibodies or T cells may thus play a direct role promoting adipogenesis, glycosaminoglycan synthesis and expression of immunomodulatory proteins within the orbits. Once the net effect of these changes has come to increase the volume of the fatty connective tissues within the orbit, then proptosis, extraocular muscle dysfunction, and periorbital congestion will ensue.

Simple satellite orbit propagator

NASA Astrophysics Data System (ADS)

Gurfil, P.

2008-06-01

An increasing number of space missions require on-board autonomous orbit determination. The purpose of this paper is to develop a simple orbit propagator (SOP) for such missions. Since most satellites are limited by the available processing power, it is important to develop an orbit propagator that will use limited computational and memory resources. In this work, we show how to choose state variables for propagation using the simplest numerical integration scheme available-the explicit Euler integrator. The new state variables are derived by the following rationale: Apply a variation-of-parameters not on the gravity-affected orbit, but rather on the gravity-free orbit, and teart the gravity as a generalized force. This ultimately leads to a state vector comprising the inertial velocity and a modified position vector, wherein the product of velocity and time is subtracted from the inertial position. It is shown that the explicit Euler integrator, applied on the new state variables, becomes a symplectic integrator, preserving the Hamiltonian and the angular momentum (or a component thereof in the case of oblateness perturbations). The main application of the proposed propagator is estimation of mean orbital elements. It is shown that the SOP is capable of estimating the mean elements with an accuracy that is comparable to a high-order integrator that consumes an order-of-magnitude more computational time than the SOP.

Bernese advances towards a global analysis of Lunar geodesy

NASA Astrophysics Data System (ADS)

Bertone, S.; Girardin, V.; Bourgoin, A.; Arnold, D.; Jaeggi, A.

2017-12-01

In this presentation we discuss our latest GRAIL-based lunar gravity fields generated with the Celestial Mechanics Approach using the planetary extension of the Bernese GNSS Software (BSW) developed at the Astronomical Institute of the University of Bern (AIUB).Based on one-way X band and two-way S-band Doppler data, we perform orbit determination by solving six initial orbital elements, dynamical parameters, and stochastic parameters in daily arcs using a least-squares adjustment. Significative improvements to our solutions come from the recent implementation of an accurate modeling of non-gravitational forces, including accelerations due to solar and planetary (albedo and IR) radiation pressure, based on the 28-plate macromodel to represent the GRAIL satellites. Also, as suggested in previous works, we deal with imperfections in the modeling of solar eclipses by both an accurate data screening at mid-latitudes and by taking into account solar panel voltage data in our processing. Empirical and pseudo-stochastic parameters are estimated on top of our dynamical modeling to absorb its deficiencies. We analyze the impact of different parametrizations using either pulses (i.e., instantaneous velocity changes) and piecewise constant accelerations (PCA) on our orbits.Based on these improved orbits, one- and two-way Doppler and KBRR data are then used together with an appropriate weighting for a combined orbit and gravity field determination process.We present our latest solutions of the lunar gravity field, based on the recent GRAIL GRGM900C gravity field (as validation of our modeling and parametrization) and on iterations from the SELENE SGM150J gravity field (to check the independence of our solution). We detail our procedure to gradually enlarge the parameter space while adding new data to our gravity field solution. In addition, we present our latest solution for the Moon tidal Love number k_2.Moreover, some important lunar geophysical parameters are best obtained by processing data from Lunar Laser Ranging (LLR) stations. We analyze the impact of a combined processing on the recovery of a set of parameters and we provide some preliminary results.Finally, we compare all of our results with the most recent solutions of the lunar gravity field and of other geodetic parameters released by other groups.

Applying Parallel Processing Techniques to Tether Dynamics Simulation

NASA Technical Reports Server (NTRS)

Wells, B. Earl

1996-01-01

The focus of this research has been to determine the effectiveness of applying parallel processing techniques to a sizable real-world problem, the simulation of the dynamics associated with a tether which connects two objects in low earth orbit, and to explore the degree to which the parallelization process can be automated through the creation of new software tools. The goal has been to utilize this specific application problem as a base to develop more generally applicable techniques.

Independent Orbiter Assessment (IOA): Analysis of the communication and tracking subsystem

NASA Technical Reports Server (NTRS)

Gardner, J. R.; Robinson, W. M.; Trahan, W. H.; Daley, E. S.; Long, W. C.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Communication and Tracking hardware. The IOA analysis process utilized available Communication and Tracking hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Dynamics of binary and planetary-system interaction with disks - Eccentricity changes

NASA Technical Reports Server (NTRS)

Atrymowicz, Pawel

1992-01-01

Protostellar and protoplanetary systems, as well as merging galactic nuclei, often interact tidally and resonantly with the astrophysical disks via gravity. Underlying our understanding of the formation processes of stars, planets, and some galaxies is a dynamical theory of such interactions. Its main goals are to determine the geometry of the binary-disk system and, through the torque calculations, the rate of change of orbital elements of the components. We present some recent developments in this field concentrating on eccentricity driving mechanisms in protoplanetary and protobinary systems. In those two types of systems the result of the interaction is opposite. A small body embedded in a disk suffers a decrease of orbital eccentricity, whereas newly formed binary stars surrounded by protostellar disks may undergo a significant orbital evolution increasing their eccentricities.

The capture of lunar materials in low lunar orbit

NASA Technical Reports Server (NTRS)

Floyd, M. A.

1981-01-01

A scenario is presented for the retrieval of lunar materials sent into lunar orbit to be used as raw materials in space manufacturing operations. The proposal is based on the launch of material from the lunar surface by an electromagnetic mass driver and the capture of this material in low lunar orbit by a fleet of mass catchers which ferry the material to processing facilities when full. Material trajectories are analyzed using the two-body equations of motion, and intercept requirements and the sensitivity of the system to launch errors are determined. The present scenario is shown to be superior to scenarios that place a single mass catcher at the L2 libration point due to increased operations flexibility, decreased mass driver performance requirements and centralized catcher servicing.

Breakthrough in orbit determination of a binary. - In expectation of astrometric observations with high precision such as VERA and JASMINE -

NASA Astrophysics Data System (ADS)

Asada, Hideki

2006-11-01

There exists a very classical inverse problem regarding orbit determination of a binary system: "when an orbital plane of two bodies is inclined with respect to the line of sight, observables are their positions projected onto a celestial sphere. How do we determine the orbital elements from observations?" A "complete exact solution" has been found. It is reviewed with some related topics.

Filling of orbital fluid management systems

NASA Technical Reports Server (NTRS)

Merino, F.; Blatt, M. H.; Thies, N. C.

1978-01-01

A study was performed with three objectives: (1) analyze fluid management system fill under orbital conditions; (2) determine what experimentation is needed; and (3) develop an experimental program. The fluid management system was a 1.06m (41.7 in) diameter pressure vessel with screen channel device. Analyses were conducted using liquid hydrogen and N2O4. The influence of helium and autogenous pressurization systems was considered. Analyses showed that fluid management system fill will be more difficult with a cryogen than with an earth storable. The key to a successful fill with cryogens is in devising techniques for filling without vent liquid, and removing trapped vapor from the screen device at tank fill completion. This will be accomplished with prechill, fill, and vapor condensation processes. Refill will require a vent and purge process, to dilute the residual helium, prior to introducing liquid. Neither prechill, chill, nor purge processes will be required for earth storables.

SPECTROSCOPIC ORBITS FOR 15 LATE-TYPE STARS

DOE Office of Scientific and Technical Information (OSTI.GOV)

Willmarth, Daryl W.; Abt, Helmut A.; Fekel, Francis C.

2016-08-01

Spectroscopic orbital elements are determined for 15 stars with periods from 8 to 6528 days with six orbits computed for the first time. Improved astrometric orbits are computed for two stars and one new orbit is derived. Visual orbits were previously determined for four stars, four stars are members of multiple systems, and five stars have Hipparcos “G” designations or have been resolved by speckle interferometry. For the nine binaries with previous spectroscopic orbits, we determine improved or comparable elements. For HD 28271 and HD 200790, our spectroscopic results support the conclusions of previous authors that the large values of their massmore » functions and lack of detectable secondary spectrum argue for the secondary in each case being a pair of low-mass dwarfs. The orbits given here may be useful in combination with future interferometric and Gaia satellite observations.« less

Geospace Magnetospheric Dynamics Mission

NASA Technical Reports Server (NTRS)

Russell, C. T.; Kluever, C.; Burch, J. L.; Fennell, J. F.; Hack, K.; Hillard, G. B.; Kurth, W. S.; Lopez, R. E.; Luhmann, J. G.; Martin, J. B.;

CALIBRATION OF EQUILIBRIUM TIDE THEORY FOR EXTRASOLAR PLANET SYSTEMS. II

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hansen, Brad M. S., E-mail: hansen@astro.ucla.edu

2012-09-20

We present a new empirical calibration of equilibrium tidal theory for extrasolar planet systems, extending a prior study by incorporating detailed physical models for the internal structure of planets and host stars. The resulting strength of the stellar tide produces a coupling that is strong enough to reorient the spins of some host stars without causing catastrophic orbital evolution, thereby potentially explaining the observed trend in alignment between stellar spin and planetary orbital angular momentum. By isolating the sample whose spins should not have been altered in this model, we also show evidence for two different processes that contribute tomore » the population of planets with short orbital periods. We apply our results to estimate the remaining lifetimes for short-period planets, examine the survival of planets around evolving stars, and determine the limits for circularization of planets with highly eccentric orbits. Our analysis suggests that the survival of circularized planets is strongly affected by the amount of heat dissipated, which is often large enough to lead to runaway orbital inflation and Roche lobe overflow.« less

Evaluation of Landsat-4 orbit determination accuracy using batch least-squares and sequential methods

NASA Astrophysics Data System (ADS)

Oza, D. H.; Jones, T. L.; Feiertag, R.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite (TDRS) System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the May 18-24, 1992, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. During this period, there were two separate orbit-adjust maneuvers on one of the TDRSS spacecraft (TDRS-East) and one small orbit-adjust maneuver for Landsat-4. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 30 meters after the filter had reached steady state.

Evaluation of Landsat-4 orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Feiertag, R.; Samii, M. V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1993-01-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite (TDRS) System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the May 18-24, 1992, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. During this period, there were two separate orbit-adjust maneuvers on one of the TDRSS spacecraft (TDRS-East) and one small orbit-adjust maneuver for Landsat-4. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 30 meters after the filter had reached steady state.

Further developments in orbit ephemeris derived neutral density

NASA Astrophysics Data System (ADS)

Locke, Travis

There are a number of non-conservative forces acting on a satellite in low Earth orbit. The one which is the most dominant and also contains the most uncertainty is atmospheric drag. Atmospheric drag is directly proportional to atmospheric density, and the existing atmospheric density models do not accurately model the variations in atmospheric density. In this research, precision orbit ephemerides (POE) are used as input measurements in an optimal orbit determination scheme in order to estimate corrections to existing atmospheric density models. These estimated corrections improve the estimates of the drag experienced by a satellite and therefore provide an improvement in orbit determination and prediction as well as a better overall understanding of the Earth's upper atmosphere. The optimal orbit determination scheme used in this work includes using POE data as measurements in a sequential filter/smoother process using the Orbit Determination Tool Kit (ODTK) software. The POE derived density estimates are validated by comparing them with the densities derived from accelerometers on board the Challenging Minisatellite Payload (CHAMP) and the Gravity Recovery and Climate Experiment (GRACE). These accelerometer derived density data sets for both CHAMP and GRACE are available from Sean Bruinsma of the Centre National d'Etudes Spatiales (CNES). The trend in the variation of atmospheric density is compared quantitatively by calculating the cross correlation (CC) between the POE derived density values and the accelerometer derived density values while the magnitudes of the two data sets are compared by calculating the root mean square (RMS) values between the two. There are certain high frequency density variations that are observed in the accelerometer derived density data but not in the POE derived density data or any of the baseline density models. These high frequency density variations are typically small in magnitude compared to the overall day-night variation. However during certain time periods, such as when the satellite is near the terminator, the variations are on the same order of magnitude as the diurnal variations. These variations can also be especially prevalent during geomagnetic storms and near the polar cusps. One of the goals of this work is to see what affect these unmodeled high frequency variations have on orbit propagation. In order to see this effect, the orbits of CHAMP and GRACE are propagated during certain time periods using different sources of density data as input measurements (accelerometer, POE, HASDM, and Jacchia 1971). The resulting orbit propagations are all compared to the propagation using the accelerometer derived density data which is used as truth. The RMS and the maximum difference between the different propagations are analyzed in order to see what effect the unmodeled density variations have on orbit propagation. These results are also binned by solar and geomagnetic activity level. The primary input into the orbit determination scheme used to produce the POE derived density estimates is a precision orbit ephemeris file. This file contains position and velocity in-formation for the satellite based on GPS and SLR measurements. The values contained in these files are estimated values and therefore contain some level of error, typically thought to be around the 5-10 cm level. The other primary focus of this work is to evaluate the effect of adding different levels of noise (0.1 m, 0.5 m, 1 m, 10 m, and 100 m) to this raw ephemeris data file before it is input into the orbit determination scheme. The resulting POE derived density estimates for each level of noise are then compared with the accelerometer derived densities by computing the CC and RMS values between the data sets. These results are also binned by solar and geomagnetic activity level.

Short arc orbit determination and Gaia alerts

NASA Astrophysics Data System (ADS)

Spoto, Federica; Tanga, Paolo; Del Vigna, Alessio; Carry, Benoit; Thuillot, William; David, Pedro; Mignard, Francois; Milani, Andrea; Tommei, Giacomo

2017-10-01

Since October 2016, the short term (ST) processing of Solar System Objects (SSOs) by Gaia is up and running, and it has produced almost 600 alerts. A crucial point in the chain is the possibility of performing a short arc orbit determination as soon as the object has been detected, which allows the follow up of the object from the ground.The method we present has been recentely developed for two mainreasons: 1) search for imminent impactors within the NEO - Confirmation Page(imminent impactors are asteroids that could impact the Earth infew days from their discovery) 2) validation of the SSO-ST Gaia pipeline.We show some good confirmations on objects that could have been discovered by Gaia, and some properties of the Gaia astrometry for the short term.

Incorporation of star measurements for the determination of orbit and attitude parameters of a geosynchronous satellite: An iterative application of linear regression

NASA Astrophysics Data System (ADS)

Phillips, D.

1980-10-01

Currently on NOAA/NESS's VIRGS system at the World Weather Building star images are being ingested on a daily basis. The image coordinates of the star locations are measured and stored. Subsequently, the information is used to determine the attitude, the misalignment angles between the spin axis and the principal axis of the satellite, and the precession rate and direction. This is done for both the 'East' and 'West' operational geosynchronous satellites. This orientation information is then combined with image measurements of earth based landmarks to determine the orbit of each satellite. The method for determining the orbit is simple. For each landmark measurement one determines a nominal position vector for the satellite by extending a ray from the landmark's position towards the satellite and intersecting the ray with a sphere with center coinciding with the Earth's center and with radius equal to the nominal height for a geosynchronous satellite. The apparent motion of the satellite around the Earth's center is then approximated with a Keplerian model. In turn the variations of the satellite's height, as a function of time found by using this model, are used to redetermine the successive satellite positions by again using the Earth based landmark measurements and intersecting rays from these landmarks with the newly determined spheres. This process is performed iteratively until convergence is achieved. Only three iterations are required.

Orbit determination error analysis and comparison of station-keeping costs for Lissajous and halo-type libration point orbits and sensitivity analysis using experimental design techniques

NASA Technical Reports Server (NTRS)

Gordon, Steven C.

1993-01-01

Spacecraft in orbit near libration point L1 in the Sun-Earth system are excellent platforms for research concerning solar effects on the terrestrial environment. One spacecraft mission launched in 1978 used an L1 orbit for nearly 4 years, and future L1 orbital missions are also being planned. Orbit determination and station-keeping are, however, required for these orbits. In particular, orbit determination error analysis may be used to compute the state uncertainty after a predetermined tracking period; the predicted state uncertainty levels then will impact the control costs computed in station-keeping simulations. Error sources, such as solar radiation pressure and planetary mass uncertainties, are also incorporated. For future missions, there may be some flexibility in the type and size of the spacecraft's nominal trajectory, but different orbits may produce varying error analysis and station-keeping results. The nominal path, for instance, can be (nearly) periodic or distinctly quasi-periodic. A periodic 'halo' orbit may be constructed to be significantly larger than a quasi-periodic 'Lissajous' path; both may meet mission requirements, but perhaps the required control costs for these orbits are probably different. Also for this spacecraft tracking and control simulation problem, experimental design methods can be used to determine the most significant uncertainties. That is, these methods can determine the error sources in the tracking and control problem that most impact the control cost (output); it also produces an equation that gives the approximate functional relationship between the error inputs and the output.

Application of the Constrained Admissible Region Multiple Hypothesis Filter to Initial Orbit Determination of a Break-up

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.

Constellation Coverage Analysis

NASA Technical Reports Server (NTRS)

Lo, Martin W. (Compiler)

1997-01-01

The design of satellite constellations requires an understanding of the dynamic global coverage provided by the constellations. Even for a small constellation with a simple circular orbit propagator, the combinatorial nature of the analysis frequently renders the problem intractable. Particularly for the initial design phase where the orbital parameters are still fluid and undetermined, the coverage information is crucial to evaluate the performance of the constellation design. We have developed a fast and simple algorithm for determining the global constellation coverage dynamically using image processing techniques. This approach provides a fast, powerful and simple method for the analysis of global constellation coverage.

Orbit Determination Error Analysis Results for the Triana Sun-Earth L2 Libration Point Mission

NASA Technical Reports Server (NTRS)

Marr, G.

2003-01-01

Using the NASA Goddard Space Flight Center's Orbit Determination Error Analysis System (ODEAS), orbit determination error analysis results are presented for all phases of the Triana Sun-Earth L1 libration point mission and for the science data collection phase of a future Sun-Earth L2 libration point mission. The Triana spacecraft was nominally to be released by the Space Shuttle in a low Earth orbit, and this analysis focuses on that scenario. From the release orbit a transfer trajectory insertion (TTI) maneuver performed using a solid stage would increase the velocity be approximately 3.1 km/sec sending Triana on a direct trajectory to its mission orbit. The Triana mission orbit is a Sun-Earth L1 Lissajous orbit with a Sun-Earth-vehicle (SEV) angle between 4.0 and 15.0 degrees, which would be achieved after a Lissajous orbit insertion (LOI) maneuver at approximately launch plus 6 months. Because Triana was to be launched by the Space Shuttle, TTI could potentially occur over a 16 orbit range from low Earth orbit. This analysis was performed assuming TTI was performed from a low Earth orbit with an inclination of 28.5 degrees and assuming support from a combination of three Deep Space Network (DSN) stations, Goldstone, Canberra, and Madrid and four commercial Universal Space Network (USN) stations, Alaska, Hawaii, Perth, and Santiago. These ground stations would provide coherent two-way range and range rate tracking data usable for orbit determination. Larger range and range rate errors were assumed for the USN stations. Nominally, DSN support would end at TTI+144 hours assuming there were no USN problems. Post-TTI coverage for a range of TTI longitudes for a given nominal trajectory case were analyzed. The orbit determination error analysis after the first correction maneuver would be generally applicable to any libration point mission utilizing a direct trajectory.

Method of determining the orbits of the small bodies in the solar system based on an exhaustive search of orbital planes

NASA Astrophysics Data System (ADS)

Bondarenko, Yu. S.; Vavilov, D. E.; Medvedev, Yu. D.

2014-05-01

A universal method of determining the orbits of newly discovered small bodies in the Solar System using their positional observations has been developed. The proposed method suggests determining geocentric distances of a small body by means of an exhaustive search for heliocentric orbital planes and subsequent determination of the distance between the observer and the points at which the chosen plane intersects with the vectors pointing to the object. Further, the remaining orbital elements are determined using the classical Gauss method after eliminating those heliocentric distances that have a fortiori low probabilities. The obtained sets of elements are used to determine the rms between the observed and calculated positions. The sets of elements with the least rms are considered to be most probable for newly discovered small bodies. Afterwards, these elements are improved using the differential method.

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.

Tracking and Data Relay Satellite (TDRS) Orbit Estimation Using an Extended Kalman Filter

NASA Technical Reports Server (NTRS)

Ward, Douglas T.; Dang, Ket D.; Slojkowski, Steve; Blizzard, Mike; Jenkins, Greg

2007-01-01

Alternatives to the Tracking and Data Relay Satellite (TDRS) orbit estimation procedure were studied to develop a technique that both produces more reliable results and is more amenable to automation than the prior procedure. The Earth Observing System (EOS) Terra mission has TDRS ephemeris prediction 3(sigma) requirements of 75 meters in position and 5.5 millimeters per second in velocity over a 1.5-day prediction span. Meeting these requirements sometimes required reruns of the prior orbit determination (OD) process, with manual editing of tracking data to get an acceptable solution. After a study of the available alternatives, the Flight Dynamics Facility (FDF) began using the Real-Time Orbit Determination (RTOD(Registered TradeMark)) Kalman filter program for operational support of TDRSs in February 2007. This extended Kalman filter (EKF) is used for daily support, including within hours after most thrusting, to estimate the spacecraft position, velocity, and solar radiation coefficient of reflectivity (C(sub R)). The tracking data used are from the Bilateration Ranging Transponder System (BRTS), selected TDRS System (TDRSS) User satellite tracking data, and Telemetry, Tracking, and Command (TT&C) data. Degraded filter results right after maneuvers and some momentum unloads provided incentive for a hybrid OD technique. The results of combining EKF strengths with the Goddard Trajectory Determination System (GTDS) Differential Correction (DC) program batch-least-squares solutions, as recommended in a 2005 paper on the chain-bias technique, are also presented.

ERTS operations and data processing

NASA Technical Reports Server (NTRS)

Gonzales, L.; Sos, J. Y.

1974-01-01

The overall communications and data flow between the ERTS spacecraft and the ground stations and processing centers are generally described. Data from the multispectral scanner and the return beam vidicon are telemetered to a primary ground station where they are demodulated, processed, and recorded. The tapes are then transferred to the NASA Data Processing Facility (NDPF) at Goddard. Housekeeping data are relayed from the prime ground stations to the Operations Control Center at Goddard. Tracking data are processed at the ground stations, and the calculated parameters are transmitted by teletype to the orbit determination group at Goddard. The ERTS orbit has been designed so that the same swaths of the ground coverage pattern viewed during one 18-day coverage cycle are repeated by the swaths viewed on all subsequent cycles. The Operations Control Center is the focal point for all communications with the spacecraft. NDPF is a job-oriented facility which processes and stores all sensor data, and which disseminates large quantities of these data to users in the form of films, computer-compatible tapes, and data collection system data.

Laser Ranging to the Lunar Reconnaissance Orbiter: improved timing and orbits

NASA Astrophysics Data System (ADS)

Mao, D.; Mcgarry, J.; Sun, X.; Torrence, M. H.; Skillman, D.; Hoffman, E.; Mazarico, E.; Rowlands, D. D.; Golder, J.; Barker, M. K.; Neumann, G. A.; Smith, D. E.; Zuber, M. T.

2013-12-01

The Laser ranging (LR) experiment to the Lunar Reconnaissance Orbiter (LRO) has been under operation for more than 4 years, since the launch of the spacecraft in June 2009. Led by NASA's Next Generation Satellite Laser Ranging(NGSLR) station at Greenbelt, Maryland, ten laser ranging stations over the world have been participating in the experiment and have collected over 3,200 hours of ranging data. These range measurements are used to monitor the behavior of the LRO clock and to generate orbital solutions for LRO. To achieve high-quality results in range, ground stations like NGSLR are using H-maser clocks to obtain a stable and continuous time baseline for the orbit solutions. An All-View GPS receiver was included at NGSLR since January 2013 which monitors the H-maser time against the master clock at the United State Naval Observatory (USNO) via the GPS satellites. NGSLR has successfully established nano-second level epoch time accuracy and 10-15 clock stability since then. Time transfer experiments using LRO as a common receiver have been verified in ground testing between NGSLR and MOBLAS7 via a ground terminal with a Lunar Orbiter Laser Altimeter (LOLA)-like receiver at Greenbelt, Maryland. Two hour-long ground tests using a LOLA-like detector and two different ground targets yielded results consistent with each other, and those from the previous 10-minute test completed one year ago. Time transfer tests between NGSLR and MOBLAS7 via LRO are ongoing. More time transfer tests are being planned from NGSLR to McDonald Laser Ranging Station (MLRS) in Texas and later from NGSLR to European satellite laser ranging (SLR) stations. Upon the completion of these time transfer experiments, nanosecond-level epoch time accuracy will be brought to stations besides NGSLR, and such high precision of the ground time can contribute to the LRO precision orbit determination (POD) process. Presently, by using the high-resolution GRAIL gravity models, the LRO orbits determined from LR data alone have a total position error of 10 meters in average, and show the same quality as those generated using conventional radiometric tracking data. In these LR orbital solutions, a bias was adjusted to compensate both the ground and spacecraft clock characteristics. By taking advantage of the knowledge we have gained through LR of the long-term stability of the LRO clock, the spacecraft clock behavior is separated from the ground station clocks and modeled over a 10-month time span in our current POD process. Here we present the results from this new approach, and further improvements in the quality of the orbital reconstruction.

Impact of GNSS orbit modeling on LEO orbit and gravity field determination

NASA Astrophysics Data System (ADS)

Arnold, Daniel; Meyer, Ulrich; Sušnik, Andreja; Dach, Rolf; Jäggi, Adrian

2017-04-01

On January 4, 2015 the Center for Orbit Determination in Europe (CODE) changed the solar radiation pressure modeling for GNSS satellites to an updated version of the empirical CODE orbit model (ECOM). Furthermore, since September 2012 CODE operationally computes satellite clock corrections not only for the 3-day long-arc solutions, but also for the non-overlapping 1-day GNSS orbits. This provides different sets of GNSS products for Precise Point Positioning, as employed, e.g., in the GNSS-based precise orbit determination of low Earth orbiters (LEOs) and the subsequent Earth gravity field recovery from kinematic LEO orbits. While the impact of the mentioned changes in orbit modeling and solution strategy on the GNSS orbits and geophysical parameters was studied in detail, their implications on the LEO orbits were not yet analyzed. We discuss the impact of the update of the ECOM and the influence of 1-day and 3-day GNSS orbit solutions on zero-difference LEO orbit and gravity field determination, where the GNSS orbits and clock corrections, as well as the Earth rotation parameters are introduced as fixed external products. Several years of kinematic and reduced-dynamic orbits for the two GRACE LEOs are computed with GNSS products based on both the old and the updated ECOM, as well as with 1- and 3-day GNSS products. The GRACE orbits are compared by means of standard validation measures. Furthermore, monthly and long-term GPS-only and combined GPS/K-band gravity field solutions are derived from the different sets of kinematic LEO orbits. GPS-only fields are validated by comparison to combined GPS/K-band solutions, while the combined solutions are validated by analysis of the formal errors, as well as by comparing them to the combined GRACE solutions of the European Gravity Service for Improved Emergency Management (EGSIEM) project.

Micrometeoroid and Orbital Debris Threat Assessment: Mars Sample Return Earth Entry Vehicle

NASA Technical Reports Server (NTRS)

Christiansen, Eric L.; Hyde, James L.; Bjorkman, Michael D.; Hoffman, Kevin D.; Lear, Dana M.; Prior, Thomas G.

2011-01-01

This report provides results of a Micrometeoroid and Orbital Debris (MMOD) risk assessment of the Mars Sample Return Earth Entry Vehicle (MSR EEV). The assessment was performed using standard risk assessment methodology illustrated in Figure 1-1. Central to the process is the Bumper risk assessment code (Figure 1-2), which calculates the critical penetration risk based on geometry, shielding configurations and flight parameters. The assessment process begins by building a finite element model (FEM) of the spacecraft, which defines the size and shape of the spacecraft as well as the locations of the various shielding configurations. This model is built using the NX I-deas software package from Siemens PLM Software. The FEM is constructed using triangular and quadrilateral elements that define the outer shell of the spacecraft. Bumper-II uses the model file to determine the geometry of the spacecraft for the analysis. The next step of the process is to identify the ballistic limit characteristics for the various shield types. These ballistic limits define the critical size particle that will penetrate a shield at a given impact angle and impact velocity. When the finite element model is built, each individual element is assigned a property identifier (PID) to act as an index for its shielding properties. Using the ballistic limit equations (BLEs) built into the Bumper-II code, the shield characteristics are defined for each and every PID in the model. The final stage of the analysis is to determine the probability of no penetration (PNP) on the spacecraft. This is done using the micrometeoroid and orbital debris environment definitions that are built into the Bumper-II code. These engineering models take into account orbit inclination, altitude, attitude and analysis date in order to predict an impacting particle flux on the spacecraft. Using the geometry and shielding characteristics previously defined for the spacecraft and combining that information with the environment model calculations, the Bumper-II code calculates a probability of no penetration for the spacecraft.

GNSS Radio Occultation Excess Phase Processing with Integrated Uncertainty Estimation for Thermodynamic Cal/Val of Passive Atmospheric Sounders and Climate Science

NASA Astrophysics Data System (ADS)

Innerkofler, J.; Pock, C.; Kirchengast, G.; Schwaerz, M.; Jaeggi, A.; Andres, Y.; Marquardt, C.; Hunt, D.; Schreiner, W. S.; Schwarz, J.

2017-12-01

Global Navigation Satellite System (GNSS) radio occultation (RO) is a highly valuable satellite remote sensing technique for atmospheric and climate sciences, including calibration and validation (cal/val) of passive sounding instruments such as radiometers. It is providing accurate and precise measurements in the troposphere and stratosphere regions with global coverage, long-term stability, and virtually all-weather capability since 2001. For fully exploiting the potential of RO data as a cal/val reference and climate data record, uncertainties attributed to the data need to be assessed. Here we focus on the atmospheric excess phase data, based on the raw occultation tracking and orbit data, and its integrated uncertainty estimation within the new Reference Occultation Processing System (rOPS) developed at the WEGC. These excess phases correspond to integrated refractivity, proportional to pressure/temperature and water vapor, and are therefore highly valuable reference data for thermodynamic cal/val of passive (radiometric) sounder data. In order to enable high accuracy of the excess phase profiles, accurate orbit positions and velocities as well as clock estimates of the GNSS transmitter satellites and RO receiver satellites are determined using the Bernese and Napeos orbit determination software packages. We find orbit uncertainty estimates of about 5 cm (position) / 0.05 mm/s (velocity) for daily orbits for the MetOp, GRACE, and CHAMP RO missions, and decreased uncertainty estimates near 20 cm (position) / 0.2 mm/s (velocity) for the COSMIC RO mission. The strict evaluation and quality control of the position, velocity, and clock accuracies of the daily LEO and GNSS orbits assure smallest achievable uncertainties in the excess phase data. We compared the excess phase profiles from WEGC against profiles from EUMETSAT and UCAR. Results show good agreement in line with the estimated uncertainties, with millimetric differences in the upper stratosphere and mesosphere and centimetric differences in the troposphere, where the excess phases amount to beyond 100 m. This underlines the potential for a new fundamental cal/val reference and climate data record based on atmospheric excess phases from RO, given their narrow uncertainty and independence from background data.

Flight Test Results from the Low Power Transceiver Communications and Navigation Demonstration on Shuttle (CANDOS)

NASA Technical Reports Server (NTRS)

Rush, John; Israel, David; Harlacher, Marc; Haas, Lin

2003-01-01

The Low Power Transceiver (LPT) is an advanced signal processing platform that offers a configurable and reprogrammable capability for supporting communications, navigation and sensor functions for mission applications ranging from spacecraft TT&C and autonomous orbit determination to sophisticated networks that use crosslinks to support communications and real-time relative navigation for formation flying. The LPT is the result of extensive collaborative research under NASNGSFC s Advanced Technology Program and ITT Industries internal research and development efforts. Its modular, multi-channel design currently enables transmitting and receiving communication signals on L- or S-band frequencies and processing GPS L-band signals for precision navigation. The LPT flew as a part of the GSFC Hitchhiker payload named Fast Reaction Experiments Enabling Science Technology And Research (FREESTAR) on-board Space Shuttle Columbia s final mission. The experiment demonstrated functionality in GPS-based navigation and orbit determination, NASA STDN Ground Network communications, space relay communications via the NASA TDRSS, on-orbit reconfiguration of the software radio, the use of the Internet Protocol (IP) for TT&C, and communication concepts for space based range safety. All data from the experiment was recovered and, as a result, all primary and secondary objectives of the experiment were successful. This paper presents the results of the LPTs maiden space flight as a part of STS- 107.

On a possible additional component in an eclipsing binary system HS 2231 + 2441

NASA Astrophysics Data System (ADS)

Vidmachenko, A. P.; Romanyuk, Ya. O.; Shliahetskaya, Ya. O.

2016-05-01

Timing method based on the registration period of variations of a periodic process, associated with the star. The study of stellar eclipsing binary system for a long time allows a series of several transits, depending on the orbital period of the satellite smaller. We present a photometric study of system of the type HW Vir HS 2231 + 2441. Photometric data processing was performed using C-MuniWin Version 1.2.30 program. The accuracy of values for each observation point is in the range 0,003...0,009m for different nights. The calculated ephemeris determined from the light curve by fitting of arc of minimums to the nuclei of primary and secondary eclipses. The amplitude of the periodic changes of minimums moments that arise from the orbital motion of a close pair of stars around the barycenter of the triple system, is less than 0.0008 days (1.15 minutes). It was found that the periodic variation of the orbital period can be explained by the gravitational influence of a third companion on the central binary system with an orbital period of about 97±10d. Periodogram analysis of the observational data series indicate also on the periodicity with values of 48±5d and 195±15d, but with substantially less reliably

The Magnetospheric Multiscale Mission...Resolving Fundamental Processes in Space Plasmas

NASA Technical Reports Server (NTRS)

Curtis, S.

1999-01-01

The Magnetospheric Multiscale (MMS) mission is a multiple-spacecraft Solar-Terrestrial Probe designed to study the microphysics of magnetic reconnection, charged particle acceleration, and turbulence in key boundary regions of Earth's magnetosphere. These three processes, which control the flow of energy, mass, and momentum within and across plasma boundaries, occur throughout the universe and are fundamental to our understanding of astrophysical and solar system plasmas. Only in Earth's magnetosphere, however, are they readily accessible for sustained study through in-situ measurement. MMS will employ five co-orbiting spacecraft identically instrumented to measure electric and magnetic fields, plasmas, and energetic particles. The initial parameters of the individual spacecraft orbits will be designed so that the spacecraft formation will evolve into a three-dimensional configuration near apogee, allowing MMS to differentiate between spatial and temporal effects and to determine the three dimensional geometry of plasma, field, and current structures. In order to sample all of the magnetospheric boundary regions, MMS will employ a unique four-phase orbital strategy involving carefully sequenced changes in the local time and radial distance of apogee and, in the third phase, a change in orbit inclination from 10 degrees to 90 degrees. The nominal mission operational lifetime is two years. Launch is currently scheduled for 2006.

A Novel Method for Precise Onboard Real-Time Orbit Determination with a Standalone GPS Receiver

PubMed Central

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

Marking Tests to Certify Part Identification Processes for Use in Low Earth Orbit

NASA Technical Reports Server (NTRS)

Roxby, D. L.

2015-01-01

The primary purpose for the MISSE marking tests was to define Data Matrix symbol marking processes that will remain readable after exposure to Low Earth Orbit environments. A wide range of different Data Matrix symbol marking processes and materials, including some still under development, were evaluated. The samples flown on MISSE 1 and 2 were in orbit for 3 years and 348 days, MISSE 3 and 4 were in orbit for 1 year and 15 days, MISSE 6 was in orbit for 1 year and 130 days, and MISSE 8 was in orbit for 2 years and 55 days. The initial MISSE marking tests clearly reflected that intrusive marking processes can be successfully used for this purpose. All of the intrusive marking processes tested exceeded program expectations and met 100 percent of the principle investigators objectives. However, subsequent tests demonstrated that some additive marking processes will also satisfy the requirements. This was an unexpected result.

Direct Imaging Of Long Period Radial Velocity Targets With NICI

NASA Astrophysics Data System (ADS)

Salter, Graeme S.; Tinney, Chris G.; Wittenmyer, Robert A.; Jenkins, James S.; Jones, Hugh R. A.; O'Toole, Simon J.

2014-01-01

We are finally entering an era where radial velocity and direct imaging parameter spaces are starting to overlap. Radial velocity measurements provide us with a minimum mass for an orbiting companion (the mass as a function of the inclination of the system). By following up these long period radial velocity detections with direct imaging we can determine whether a trend seen is due to an orbiting planet at low inclination or an orbiting brown dwarf at high inclination. In the event of a non-detection we are still able to put a limit on the maximum mass of the orbiting body. The Anglo-Australian Planet Search is one of the longest baseline radial velocity planet searches in existence, amongst its targets are many that show long period trends in the data. Here we present our direct imaging survey of these objects with our results to date. ADI Observations have been made using NICI (Near Infrared Coronagraphic Imager) on Gemini South and analysed using an in house, LOCI-like, post processing.

Independent Orbiter Assessment (IOA): FMEA/CIL assessment

NASA Technical Reports Server (NTRS)

Saiidi, Mo J.; Swain, L. J.; Compton, J. M.

1988-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. Direction was given by the Orbiter and GFE Projects Office to perform the hardware analysis and assessment using the instructions and ground rules defined in NSTS 22206. The IOA analysis features a top-down approach to determine hardware failure modes, criticality, and potential critical items. To preserve independence, the anlaysis was accomplished without reliance upon the results contained within the NASA and prime contractor FMEA/CIL documentation. The assessment process compares the independently derived failure modes and criticality assignments to the proposed NASA Post 51-L FMEA/CIL documentation. When possible, assessment issues are discussed and resolved with the NASA subsystem managers. The assessment results for each subsystem are summarized. The most important Orbiter assessment finding was the previously unknown stuck autopilot push-button criticality 1/1 failure mode, having a worst case effect of loss of crew/vehicle when a microwave landing system is not active.

Development of flight experiment work performance and workstation interface requirements, part 1. Technical report and appendices A through G

NASA Technical Reports Server (NTRS)

Hatterick, R. G.

1973-01-01

A skill requirement definition method was applied to the problem of determining, at an early stage in system/mission definition, the skills required of on-orbit crew personnel whose activities will be related to the conduct or support of earth-orbital research. The experiment data base was selected from proposed experiments in NASA's earth orbital research and application investigation program as related to space shuttle missions, specifically those being considered for Sortie Lab. Concepts for two integrated workstation consoles for Sortie Lab experiment operations were developed, one each for earth observations and materials sciences payloads, utilizing a common supporting subsystems core console. A comprehensive data base of crew functions, operating environments, task dependencies, task-skills and occupational skills applicable to a representative cross section of earth orbital research experiments is presented. All data has been coded alphanumerically to permit efficient, low cost exercise and application of the data through automatic data processing in the future.

Determination of shuttle orbiter center of gravity from flight measurements

NASA Technical Reports Server (NTRS)

Hinson, E. W.; Nicholson, J. Y.; Blanchard, R. C.

1991-01-01

Flight measurements of pitch, yaw, and roll rates and the resultant rotationally induced linear accelerations during three orbital maneuvers on Shuttle mission space transportation system (STS) 61-C were used to calculate the actual orbiter center-of-gravity location. The calculation technique reduces error due to lack of absolute calibration of the accelerometer measurements and compensates for accelerometer temperature bias and for the effects of gravity gradient. Accuracy of the technique was found to be limited by the nonrandom and asymmetrical distribution of orbiter structural vibration at the accelerometer mounting location. Fourier analysis of the vibration was performed to obtain the power spectral density profiles which show magnitudes in excess of 10(exp 4) ug (sup 2)/Hz for the actual vibration and over 500 ug (sup 2)/Hz for the filtered accelerometer measurements. The data from this analysis provide a characterization of the Shuttle acceleration environment which may be useful in future studies related to accelerometer system application and zero-g investigations or processes.

TDRSS-user orbit determination using batch least-squares and sequential methods

NASA Astrophysics Data System (ADS)

Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1993-02-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.

TDRSS-user orbit determination using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1993-01-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Gene Peavler works in the wheel area on the orbiter Discovery. The vehicle has undergone Orbiter Major Modifications in the past year. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

NASA Image and Video Library

2003-12-09

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Gene Peavler works in the wheel area on the orbiter Discovery. The vehicle has undergone Orbiter Major Modifications in the past year. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation

NASA Technical Reports Server (NTRS)

Miller, J. K.; Antreasian, P. J.; Georgini, J.; Owen, W. M.; Williams, B. G.; Yeomans, D. K.

1995-01-01

Navigation of the orbit phase of the Near Earth steroid Rendezvous (NEAR) mission will re,quire determination of certain physical parameters describing the size, shape, gravity field, attitude and inertial properties of Eros. Prior to launch, little was known about Eros except for its orbit which could be determined with high precision from ground based telescope observations. Radar bounce and light curve data provided a rough estimate of Eros shape and a fairly good estimate of the pole, prime meridian and spin rate. However, the determination of the NEAR spacecraft orbit requires a high precision model of Eros's physical parameters and the ground based data provides only marginal a priori information. Eros is the principal source of perturbations of the spacecraft's trajectory and the principal source of data for determining the orbit. The initial orbit determination strategy is therefore concerned with developing a precise model of Eros. The original plan for Eros orbital operations was to execute a series of rendezvous burns beginning on December 20,1998 and insert into a close Eros orbit in January 1999. As a result of an unplanned termination of the rendezvous burn on December 20, 1998, the NEAR spacecraft continued on its high velocity approach trajectory and passed within 3900 km of Eros on December 23, 1998. The planned rendezvous burn was delayed until January 3, 1999 which resulted in the spacecraft being placed on a trajectory that slowly returns to Eros with a subsequent delay of close Eros orbital operations until February 2001. The flyby of Eros provided a brief glimpse and allowed for a crude estimate of the pole, prime meridian and mass of Eros. More importantly for navigation, orbit determination software was executed in the landmark tracking mode to determine the spacecraft orbit and a preliminary shape and landmark data base has been obtained. The flyby also provided an opportunity to test orbit determination operational procedures that will be used in February of 2001. The initial attitude and spin rate of Eros, as well as estimates of reference landmark locations, are obtained from images of the asteroid. These initial estimates are used as a priori values for a more precise refinement of these parameters by the orbit determination software which combines optical measurements with Doppler tracking data to obtain solutions for the required parameters. As the spacecraft is maneuvered; closer to the asteroid, estimates of spacecraft state, asteroid attitude, solar pressure, landmark locations and Eros physical parameters including mass, moments of inertia and gravity harmonics are determined with increasing precision. The determination of the elements of the inertia tensor of the asteroid is critical to spacecraft orbit determination and prediction of the asteroid attitude. The moments of inertia about the principal axes are also of scientific interest since they provide some insight into the internal mass distribution. Determination of the principal axes moments of inertia will depend on observing free precession in the asteroid's attitude dynamics. Gravity harmonics are in themselves of interest to science. When compared with the asteroid shape, some insight may be obtained into Eros' internal structure. The location of the center of mass derived from the first degree harmonic coefficients give a direct indication of overall mass distribution. The second degree harmonic coefficients relate to the radial distribution of mass. Higher degree harmonics may be compared with surface features to gain additional insight into mass distribution. In this paper, estimates of Eros physical parameters obtained from the December 23,1998 flyby will be presented. This new knowledge will be applied to simplification of Eros orbital operations in February of 2001. The resulting revision to the orbit determination strategy will also be discussed.

Activities of JAXA's Innovative Technology Center on Space Debris Observation

NASA Astrophysics Data System (ADS)

Yanagisawa, T.; Kurosaki, H.; Nakajima, A.

The innovative technology research center of JAXA is developing observational technologies for GEO objects in order to cope with the space debris problem. The center had constructed the optical observational facility for space debris at Mt. Nyukasa, Nagano in 2006. As observational equipments such as CCD cameras and telescopes were set up, the normal observation started. In this paper, the detail of the facilities and its activities are introduced. The observational facility contains two telescopes and two CCD cameras. The apertures of the telescopes are 35cm and 25 cm, respectively. One CCD camera in which 2K2K chip is installed can observe a sky region of 1.3 times 1.3-degree using the 35cm telescope. The other CCD camera that contains two 4K2K chips has an ability to observe 2.6 times 2.6-degree's region with the 25cm telescope. One of our main objectives is to detect faint GEO objects that are not catalogued. Generally, the detection limit of GEO object is determined by the aperture of the telescope. However, by improving image processing techniques, the limit may become low. We are developing some image processing methods that use many CCD frames to detect faint objects. We are trying to use FPGA (Field Programmable Gate Array) system to reduce analyzing time. By applying these methods to the data taken by a large telescope, the detection limit will be significantly lowered. The orbital determination of detected GEO debris is one of the important things to do. Especially, the narrow field view of an optical telescope hinders us from re-detection of the GEO debris for the orbital determination. Long observation time is required for one GEO object for the orbital determination that is inefficient. An effective observation strategy should be considered. We are testing one observation method invented by Umehara that observes one inertia position in the space. By observing one inertia position for two nights, a GEO object that passed through the position in the first night must pass through the position in the second night. The rough orbit is determined from two nights' data. The test observation showed that this method was able to detect many GEO objects and determined their orbits by three nights' observations. We also joined the campaign observations of IADC(Inter-Agency Space Debris Coordination Committee). By analyzing the observed data with the method that we developed, 88 catalogued and 38 un-catalogued objects were detected. The magnitude of the faintest object detected in this campaign observation was 18.5. The object is un-detectable by human inspection.

Localized and Spectroscopic Orbitals: Squirrel Ears on Water.

ERIC Educational Resources Information Center

Martin, R. Bruce

1988-01-01

Reexamines the electronic structure of water considering divergent views. Discusses several aspects of molecular orbital theory using spectroscopic molecular orbitals and localized molecular orbitals. Gives examples for determining lowest energy spectroscopic orbitals. (ML)

A simplex method for the orbit determination of maneuvering satellites

NASA Astrophysics Data System (ADS)

Chen, JianRong; Li, JunFeng; Wang, XiJing; Zhu, Jun; Wang, DanNa

2018-02-01

A simplex method of orbit determination (SMOD) is presented to solve the problem of orbit determination for maneuvering satellites subject to small and continuous thrust. The objective function is established as the sum of the nth powers of the observation errors based on global positioning satellite (GPS) data. The convergence behavior of the proposed method is analyzed using a range of initial orbital parameter errors and n values to ensure the rapid and accurate convergence of the SMOD. For an uncontrolled satellite, the orbit obtained by the SMOD provides a position error compared with GPS data that is commensurate with that obtained by the least squares technique. For low Earth orbit satellite control, the precision of the acceleration produced by a small pulse thrust is less than 0.1% compared with the calibrated value. The orbit obtained by the SMOD is also compared with weak GPS data for a geostationary Earth orbit satellite over several days. The results show that the position accuracy is within 12.0 m. The working efficiency of the electric propulsion is about 67% compared with the designed value. The analyses provide the guidance for subsequent satellite control. The method is suitable for orbit determination of maneuvering satellites subject to small and continuous thrust.

Sentinel-1A - First precise orbit determination results

NASA Astrophysics Data System (ADS)

Peter, H.; Jäggi, A.; Fernández, J.; Escobar, D.; Ayuga, F.; Arnold, D.; Wermuth, M.; Hackel, S.; Otten, M.; Simons, W.; Visser, P.; Hugentobler, U.; Féménias, P.

2017-09-01

Sentinel-1A is the first satellite of the European Copernicus programme. Equipped with a Synthetic Aperture Radar (SAR) instrument the satellite was launched on April 3, 2014. Operational since October 2014 the satellite delivers valuable data for more than two years. The orbit accuracy requirements are given as 5 cm in 3D. In order to fulfill this stringent requirement the precise orbit determination (POD) is based on the dual-frequency GPS observations delivered by an eight-channel GPS receiver. The Copernicus POD (CPOD) Service is in charge of providing the orbital and auxiliary products required by the PDGS (Payload Data Ground Segment). External orbit validation is regularly performed by comparing the CPOD Service orbits to orbit solutions provided by POD expert members of the Copernicus POD Quality Working Group (QWG). The orbit comparisons revealed systematic orbit offsets mainly in radial direction (approx. 3 cm). Although no independent observation technique (e.g. DORIS, SLR) is available to validate the GPS-derived orbit solutions, comparisons between the different antenna phase center variations and different reduced-dynamic orbit determination approaches used in the various software packages helped to detect the cause of the systematic offset. An error in the given geometry information about the satellite has been found. After correction of the geometry the orbit validation shows a significant reduction of the radial offset to below 5 mm. The 5 cm orbit accuracy requirement in 3D is fulfilled according to the results of the orbit comparisons between the different orbit solutions from the QWG.

Copernicus POD Service: Ready for Sentinel-3

NASA Astrophysics Data System (ADS)

Peter, H.; Fernández, J.; Escobar, D.; Féménias, P.; Flohrer, C.; Otten, M.

2015-12-01

The Copernicus POD Service is part of the Copernicus PDGS Ground Segment of the Sentinel 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. The Sentinel-1, -2 & -3 missions have different but very demanding requirements in terms of orbital accuracy and timeliness. Orbital products in Near Real Time (latency: 30 min), Short Time Critical (1.5 days) and Non-time Critical (20-30 days) are required. The accuracy requirements are very challenging, targeting 5 cm in 3D for Sentinel-1 and 2-3 cm in radial direction for Sentinel-3. Sentinel-3A carries, in addition to a GPS receiver a laser retro reflector and a DORIS receiver. On the one hand, the three different techniques GPS, SLR and DORIS make POD more complex but, on the other hand, it is very helpful to have independent techniques available for validation of the orbit results. The successful POD processing for Sentinel-1A is a good preparation for Sentinel-3A due to the similar demanding orbit accuracy requirements. The Copernicus POD Service is ready for Sentinel-3A and the service will process GPS and SLR data routinely and has the capacity to process DORIS in NTC and reprocessing campaigns. The three independent orbit determination techniques on Sentinel-3 offer big potential for scientific exploitation. Carrying all three techniques together makes the satellite, e.g., very useful for combining all the techniques on observation level as it could only be done for Jason-2 until now. The Sentinel POD Quality Working Group strongly supporting the CPOD Service delivers additional orbit solutions to validate the CPOD results independently. The recommendations from this body guarantee that the CPOD Service is updated following state-of-the-art algorithms, models and conventions. The QWG also focuses on the scientific exploitation of the Sentinel missions. The current status of the CPOD Service is presented operating Sentinel-1A and -2A and its readiness for Sentinel-3A. It is shown how the quality and the timeliness of the products are guaranteed. Possibilities for scientific exploitation of the Sentinel-3 mission also in synergy with other Earth Observation and Sentinel missions are presented.

The location of Airy-0, the Mars prime meridian reference, from stereo photogrammetric processing of THEMIS IR imaging and digital elevation data

NASA Astrophysics Data System (ADS)

Duxbury, T. C.; Christensen, P.; Smith, D. E.; Neumann, G. A.; Kirk, R. L.; Caplinger, M. A.; Albee, A. A.; Seregina, N. V.; Neukum, G.; Archinal, B. A.

2014-12-01

The small crater Airy-0 was selected from Mariner 9 images to be the reference for the Mars prime meridian. Initial analyses in the year 2000 tied Viking Orbiter and Mars Orbiter Camera images of Airy-0 to the evolving Mars Orbiter Laser Altimeter global digital terrain model to update the location of Airy-0. Based upon this tie and radiometric tracking of landers/rovers from Earth, new expressions for the Mars spin axis direction, spin rate, and prime meridian epoch value were produced to define the orientation of the Martian surface in inertial space over time. Since the Mars Global Surveyor mission and Mars Orbiter Laser Altimeter global digital terrain model were completed some time ago, a more exhaustive study has been performed to determine the accuracy of the Airy-0 location and orientation of Mars at the standard epoch. Thermal Emission Imaging System (THEMIS) IR image cubes of the Airy and Gale crater regions were tied to the global terrain grid using precision stereo photogrammetric image processing techniques. The Airy-0 location was determined to be about 0.001° east of its predicted location using the currently defined International Astronomical Union (IAU) prime meridian location. Information on this new location and how it was derived will be provided to the NASA Mars Exploration Program Geodesy and Cartography Working Group for their assessment. This NASA group will make a recommendation to the IAU Working Group on Cartographic Coordinates and Rotational Elements to update the expression for the Mars spin axis direction, spin rate, and prime meridian location.

A time-efficient implementation of Extended Kalman Filter for sequential orbit determination and a case study for onboard application

NASA Astrophysics Data System (ADS)

Tang, Jingshi; Wang, Haihong; Chen, Qiuli; Chen, Zhonggui; Zheng, Jinjun; Cheng, Haowen; Liu, Lin

2018-07-01

Onboard orbit determination (OD) is often used in space missions, with which mission support can be partially accomplished autonomously, with less dependency on ground stations. In major Global Navigation Satellite Systems (GNSS), inter-satellite link is also an essential upgrade in the future generations. To serve for autonomous operation, sequential OD method is crucial to provide real-time or near real-time solutions. The Extended Kalman Filter (EKF) is an effective and convenient sequential estimator that is widely used in onboard application. The filter requires the solutions of state transition matrix (STM) and the process noise transition matrix, which are always obtained by numerical integration. However, numerically integrating the differential equations is a CPU intensive process and consumes a large portion of the time in EKF procedures. In this paper, we present an implementation that uses the analytical solutions of these transition matrices to replace the numerical calculations. This analytical implementation is demonstrated and verified using a fictitious constellation based on selected medium Earth orbit (MEO) and inclined Geosynchronous orbit (IGSO) satellites. We show that this implementation performs effectively and converges quickly, steadily and accurately in the presence of considerable errors in the initial values, measurements and force models. The filter is able to converge within 2-4 h of flight time in our simulation. The observation residual is consistent with simulated measurement error, which is about a few centimeters in our scenarios. Compared to results implemented with numerically integrated STM, the analytical implementation shows results with consistent accuracy, while it takes only about half the CPU time to filter a 10-day measurement series. The future possible extensions are also discussed to fit in various missions.

Discovery of orbital decay in SMC X-1

NASA Technical Reports Server (NTRS)

Levine, A.; Rappaport, S.; Boynton, P.; Deeter, J.; Nagase, F.

1992-01-01

The results are reported of three observations of the binary X ray pulsar SMC X-1 with the Ginga satellite. Timing analyses of the 0.71 s X ray pulsations yield Doppler delay curves which, in turn, provide the most accurate determination of the SMC X-1 orbital parameters available to date. The orbital phase of the 3.9 day orbit is determined in May 1987, Aug. 1988, and Aug. 1988 with accuracies of 11, 1, and 3.5 s, respectively. These phases are combined with two previous determinations of the orbital phase to yield the rate of change in the orbital period: P sub orb/P sub orb = (-3.34 + or - 0.023) x 10(exp -6)/yr. An interpretation of this measurement and the known decay rate for the orbit of Cen X-3 is made in the context of tidal evolution. Finally, a discussion is presented of the relation among the stellar evolution, orbital decay, and neutron star spinup time scales for the SMC X-1 system.

Constraints on the perturbed mutual motion in Didymos due to impact-induced deformation of its primary after the DART impact

NASA Astrophysics Data System (ADS)

Hirabayashi, M.; Schwartz, S. R.; Yu, Y.; Davis, A. B.; Chesley, S. R.; Fahnestock, E.; Michel, P.; Richardson, D. C.; Naidu, S.; Scheeres, D. J.; Cheng, A. F.; Rivkin, A.; Benner, L.

2017-12-01

(65803) Didymos is a binary near-Earth asteroid that consists of a top-shaped primary body rotating at a spin period of 2.26 hr and a secondary body orbiting around it at an orbital period of 11.92 hr. This asteroid is the target of the proposed NASA Double Asteroid Redirection Test (DART), which is part of the Asteroid Impact & Deflection Assessment (AIDA) mission concept. The goal of DART is to impact the secondary with the spacecraft and measure the momentum transfer by observing the perturbation of the orbital period of the system after the impact. Achieving this goal requires careful accounting for physical uncertainties that prevent accurate measurement of the momentum transfer. Here, we examine a scenario that might affect the momentum transfer measurement and a possible solution to avoiding issues due to this scenario. The primary's spin period is close to the spin barrier of rubble-pile asteroids, i.e., 2.3 hr. Also, some particles ejected from the secondary due to the DART impact may reach the primary and induce landslides or internal deformation of the primary, changing the gravity field. We have developed a numerical simulation technique for investigating how the mutual orbit of the system varies due to symmetric shape deformation of the primary along its spin axis after the DART impact. We find that if the deformation process occurs, the orbital period can change significantly, depending on the magnitude of the shape deformation. The mission currently plans a nearly head-on collision of the DART impactor with the secondary, making the orbital period of the system shorter. Our simulations show that since the deformation process always causes the primary to become more oblate, it shortens the orbital period as well. We also propose precise measurement of the primary's spin state to determine the deformation of the primary. This relies on the fact that any deformation process changes the spin state of the primary consistent with angular momentum conservation. Further investigations on this problem may improve the accuracy of the momentum transfer measurement for the AIDA mission.

Orbit Determination and Navigation of the Time History of Events and Macroscale Interactions during Substorms (THEMIS)

NASA Technical Reports Server (NTRS)

Morinelli, Patrick; Cosgrove, Jennifer; Blizzard, Mike; Robertson, Mike

2007-01-01

This paper provides an overview of the launch and early orbit activities performed by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of five probes comprising the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft. The FDF was tasked to support THEMIS in a limited capacity providing backup orbit determination support for validation purposes for all five THEMIS probes during launch plus 30 days in coordination with University of California Berkeley Flight Dynamics Center (UCB/FDC)2. The FDF's orbit determination responsibilities were originally planned to be as a backup to the UCB/FDC for validation purposes only. However, various challenges early on in the mission and a Spacecraft Emergency declared thirty hours after launch placed the FDF team in the role of providing the orbit solutions that enabled contact with each of the probes and the eventual termination of the Spacecraft Emergency. This paper details the challenges and various techniques used by the GSFC FDF team to successfully perform orbit determination for all five THEMIS probes during the early mission. In addition, actual THEMIS orbit determination results are presented spanning the launch and early orbit mission phase. Lastly, this paper enumerates lessons learned from the THEMIS mission, as well as demonstrates the broad range of resources and capabilities within the FDF for supporting critical launch and early orbit navigation activities, especially challenging for constellation missions.

Orbit Determination and Navigation of the Time History of Events and Macroscale Interactions during Substorms (THEMIS)

NASA Technical Reports Server (NTRS)

Morinelli, Patrick; Cosgrove, jennifer; Blizzard, Mike; Nicholson, Ann; Robertson, Mika

2007-01-01

This paper provides an overview of the launch and early orbit activities performed by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of five probes comprising the Time History of Events and Macroscale Interactions during Substorms (THEMIS) spacecraft. The FDF was tasked to support THEMIS in a limited capacity providing backup orbit determination support for validation purposes for all five THEMIS probes during launch plus 30 days in coordination with University of California Berkeley Flight Dynamics Center (UCB/FDC). The FDF's orbit determination responsibilities were originally planned to be as a backup to the UCB/FDC for validation purposes only. However, various challenges early on in the mission and a Spacecraft Emergency declared thirty hours after launch placed the FDF team in the role of providing the orbit solutions that enabled contact with each of the probes and the eventual termination of the Spacecraft Emergency. This paper details the challenges and various techniques used by the GSFC FDF team to successfully perform orbit determination for all five THEMIS probes during the early mission. In addition, actual THEMIS orbit determination results are presented spanning the launch and early orbit mission phase. Lastly, this paper enumerates lessons learned from the THEMIS mission, as well as demonstrates the broad range of resources and capabilities within the FDF for supporting critical launch and early orbit navigation activities, especially challenging for constellation missions.

Copernicus POD Service Operations

NASA Astrophysics Data System (ADS)

Fernandez, Jaime; Escobar, Diego; Ayuga, Francisco; Peter, Heike; Femenias, Pierre

2015-12-01

The Copernicus POD (Precise Orbit Determination) Service is part of the Copernicus PDGS Ground Segment of the Sentinel missions. A GMV-led consortium is operating the Copernicus POD Service (CPOD) 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 (Payload Data Ground Segment). This paper describes the CPOD Service and presents the current status operating Sentinel-1A and its readiness to support the Sentinel-2A and in particular Sentinel-3A incoming Commissioning Phases, with an especial emphasis on describing the Calibration and Validation (Cal/Val) activities to be performed during the Comm. Phase. Then, it is shown how the quality of the orbital products is guaranteed through external validation activities and the role of the Copernicus POD QWG (Quality Working Group).

Estimation and filtering techniques for high-accuracy GPS applications

NASA Technical Reports Server (NTRS)

Lichten, S. M.

1989-01-01

Techniques for determination of very precise orbits for satellites of the Global Positioning System (GPS) are currently being studied and demonstrated. These techniques can be used to make cm-accurate measurements of station locations relative to the geocenter, monitor earth orientation over timescales of hours, and provide tropospheric and clock delay calibrations during observations made with deep space radio antennas at sites where the GPS receivers have been collocated. For high-earth orbiters, meter-level knowledge of position will be available from GPS, while at low altitudes, sub-decimeter accuracy will be possible. Estimation of satellite orbits and other parameters such as ground station positions is carried out with a multi-satellite batch sequential pseudo-epoch state process noise filter. Both square-root information filtering (SRIF) and UD-factorized covariance filtering formulations are implemented in the software.

SD-4060OCPLT4 program, user's guide

NASA Technical Reports Server (NTRS)

Glazer, J.

1973-01-01

A brief description of the Orbit Comparison Plot (OCPLT4) program is presented, along with user information and a source program listing. In addition to correcting several errors that existed in the original program, this program incorporates the following new features: (1) For any satellite whose observations are processed by the Definitive Orbit Determination System (DODS), the orbital uncertainty estimates (OUE) can be obtained via appropriate card input with no major modification to the program. (2) All satellite-related information (e.g., plotter scales, cutoff limits, plotting frequencies) is user controlled via card input. (3) Not all components of OUE must be obtained. The user has the option of obtaining only the radial component if there is no need for the other two components. (4) The altitude and time graph formats are controlled by the user and are not stored for specific satellites.

Orbiter Capability for Providing Water to the International Space Station according to the Most Probable Flight Attitudes

NASA Technical Reports Server (NTRS)

Dunaway, Brian; Edeen, Marybeth

2000-01-01

Water to be generated by, delivered to, and processed by the International Space Station (ISS) is a critical Environmental Control and Life Support (ECLS) element, especially for the early ISS missions. A significant portion of the water required by the ISS shall be provided by the Shuttle Transportation System (STS) Orbiter. The balance of water generated by the Orbiter Fuel Cells (FC), minus that water consumed by the Orbiter Flash Evaporator System (FES) and crew, is available for transfer to the ISS. During later missions, crew respired and perspired water, as well as effluent water from the Orbiter LiOH canisters, will be collected as condensate and available for transfer to the ISS. Orbiter radiator performance provides the most variance in determining the amount of net Orbiter water available for transfer to the ISS. As radiator performance decreases, the dependence upon the FES (and FC water) increases for rejecting Orbiter waste heat. Generally, radiator performance decreases as the ISS assembly size increases (especially as solar arrays are added), and also as beta angle increases. ISS solar array deployment necessitates the use of models with articulating solar arrays (for Earth local-vertical attitudes), as array position dramatically affects Orbiter radiator performance. Recent developments in the relaxation of beta angle limitations have also increased the complexity and difficulty of providing water to the ISS. Other factors that may hinder the ability to transfer water are the number of empty Contingency Water Containers (CWCs) available, duration of open-hatch time, crew activity timeline, and full CWC storage capability. A parametric study has been accomplished that provides a quick-reference table for determining expected water generation rates for ISS missions 2A.2 through 7A.1. An hourly Orbiter water generation rate is reported according to a matrix that consists of: (1) (six) significant changes in ISS assembly configuration; (2) (four) beta angles (0 deg. , +37 deg., +53 deg. , and +75 deg.); (3) the (three) most representative ISS attitudes (XPOP-O, XPOP-180 and +XVV); (4) (four) Orbiter radiator configurations (both stowed, starboard deployed, port deployed, and both deployed) and (5) the (two) conditions (radiator inlet temperatures and fuel cell power) most consistent with sleep and wake periods. Those permutations of higher probability of occurrence than others have been identified. Another parametric study has been accomplished that provides a quick-reference table for determining expected water generation rates for ISS assembly complete missions. An hourly Orbiter water generation rate is reported according to a matrix that consists of: (1) (seven) beta angles (-75 deg., -60 deg., -30 deg., 0 deg., +30 deg., +60 deg., and +75 deg.); (2) the (nine) PYR angles that define the corners of the envelope; (3) (four) Orbiter radiator configurations (both stowed, starboard deployed, port deployed, and both deployed) and (4) the (two) conditions (radiator inlet temperatures and fuel cell power) most consistent with sleep and wake periods.

Solid Propulsion De-Orbiting and Re-Orbiting

NASA Astrophysics Data System (ADS)

Schonenborg, R. A. C.; Schoyer, H. F. R.

2009-03-01

With many "innovative" de-orbit systems (e.g. tethers, aero breaking, etc.) and with natural de-orbit, the place of impact of unburned spacecraft debris on Earth can not be determined accurately. The idea that satellites burn up completely upon re-entry is a common misunderstanding. To the best of our knowledge only rocket motors are capable of delivering an impulse that is high enough, to conduct a de-orbit procedure swiftly, hence to de-orbit at a specific moment that allows to predict the impact point of unburned spacecraft debris accurately in remote areas. In addition, swift de-orbiting will reduce the on-orbit time of the 'dead' satellite, which reduces the chance of the dead satellite being hit by other dead or active satellites, while spiralling down to Earth during a slow, 25 year, or more, natural de-orbit process. Furthermore the reduced on-orbit time reduces the chance that spacecraft batteries, propellant tanks or other components blow up and also reduces the time that the object requires tracking from Earth.The use of solid propellant for the de-orbiting of spacecraft is feasible. The main advantages of a solid propellant based system are the relatively high thrust and the facts that the system can be made autonomous quite easily and that the system can be very reliable. The latter is especially desirable when one wants to de-orbit old or 'dead' satellites that might not be able to rely anymore on their primary systems. The disadvantage however, is the addition of an extra system to the spacecraft as well as a (small) mass penalty. [1]This paper describes the above mentioned system and shows as well, why such a system can also be used to re-orbit spacecraft in GEO, at the end of their life to a graveyard orbit.Additionally the system is theoretically compared to an existing system, of which performance data is available.A swift market analysis is performed as well.

Orbital and Collisional Evolution of the Irregular Satellites

NASA Astrophysics Data System (ADS)

Nesvorný, David; Alvarellos, Jose L. A.; Dones, Luke; Levison, Harold F.

2003-07-01

The irregular moons of the Jovian planets are a puzzling part of the solar system inventory. Unlike regular satellites, the irregular moons revolve around planets at large distances in tilted and eccentric orbits. Their origin, which is intimately linked with the origin of the planets themselves, is yet to be explained. Here we report a study of the orbital and collisional evolution of the irregular satellites from times after their formation to the present epoch. The purpose of this study is to find out the features of the observed irregular moons that can be attributed to this evolution and separate them from signatures of the formation process. We numerically integrated ~60,000 test satellite orbits to map orbital locations that are stable on long time intervals. We found that the orbits highly inclined to the ecliptic are unstable due to the effect of the Kozai resonance, which radially stretches them so that satellites either escape from the Hill sphere, collide with massive inner moons, or impact the parent planet. We also found that prograde satellite orbits with large semimajor axes are unstable due to the effect of the evection resonance, which locks the orbit's apocenter to the apparent motion of the Sun around the parent planet. In such a resonance, the effect of solar tides on a resonant moon accumulates at each apocenter passage of the moon, which causes a radially outward drift of its orbital apocenter; once close to the Hill sphere, the moon escapes. By contrast, retrograde moons with large orbital semimajor axes are long-lived. We have developed an analytic model of the distant satellite orbits and used it to explain the results of our numerical experiments. In particular, we analytically studied the effect of the Kozai resonance. We numerically integrated the orbits of the 50 irregular moons (known by 2002 August 16) for 108 yr. All orbits were stable on this time interval and did not show any macroscopic variations that would indicate instabilities operating on longer time spans. The average orbits calculated from this experiment were then used to probe the collisional evolution of the irregular satellite systems. We found that (1) the large irregular moons must have collisionally eliminated many small irregular moons, thus shaping their population to the currently observed structures; (2) some dynamical families of satellites could have been formed by catastrophic collisions among the irregular moons; and (3) Phoebe's surface must have been heavily cratered by impacts from an extinct population of Saturnian irregular moons, much larger than the present one. We therefore suggest that the Cassini imaging of Phoebe in 2004 can be used to determine the primordial population of small irregular moons of Saturn. In such a case, we will also better understand the overall efficiency of the formation process of the irregular satellites and the physical conditions that existed during planetary formation. We discovered two dynamical families of tightly clustered orbits within the Jovian retrograde group. We believe that these two clusters may be the remnants of two collisionally disrupted bodies. We found that the entire Jovian retrograde group and the Saturnian inclination groups were not produced by single breakups, because the ejection velocities derived from the orbital structures of these groups greatly exceed values calculated by modern numerical models of collisional breakups. Taken together, the evidence presented here suggests that many properties of the irregular moons previously assigned to their formation process may have resulted from their later dynamical and collisional evolution. Finally, we have found that several irregular moons, namely, Pasiphae, Sinope, S/2001 J10, S/2000 S5, S/2000 S6, and S/2000 S3, have orbits characterized by secular resonances. The orbits of some of these moons apparently evolved by some slow dissipative process in the past and became captured in tiny resonant volumes.

Filter Tuning Using the Chi-Squared Statistic

NASA Technical Reports Server (NTRS)

Lilly-Salkowski, Tyler

2017-01-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) performs orbit determination (OD) for the Aqua and Aura satellites. Both satellites are located in low Earth orbit (LEO), and are part of what is considered the A-Train satellite constellation. Both spacecraft are currently in the science phase of their respective missions. The FDF has recently been tasked with delivering definitive covariance for each satellite.The main source of orbit determination used for these missions is the Orbit Determination Toolkit developed by Analytical Graphics Inc. (AGI). This software uses an Extended Kalman Filter (EKF) to estimate the states of both spacecraft. The filter incorporates force modelling, ground station and space network measurements to determine spacecraft states. It also generates a covariance at each measurement. This covariance can be useful for evaluating the overall performance of the tracking data measurements and the filter itself. An accurate covariance is also useful for covariance propagation which is utilized in collision avoidance operations. It is also valuable when attempting to determine if the current orbital solution will meet mission requirements in the future.This paper examines the use of the Chi-square statistic as a means of evaluating filter performance. The Chi-square statistic is calculated to determine the realism of a covariance based on the prediction accuracy and the covariance values at a given point in time. Once calculated, it is the distribution of this statistic that provides insight on the accuracy of the covariance.For the EKF to correctly calculate the covariance, error models associated with tracking data measurements must be accurately tuned. Over estimating or under estimating these error values can have detrimental effects on the overall filter performance. The filter incorporates ground station measurements, which can be tuned based on the accuracy of the individual ground stations. It also includes measurements from the NASA space network (SN), which can be affected by the assumed accuracy of the TDRS satellite state at the time of the measurement.The force modelling in the EKF is also an important factor that affects the propagation accuracy and covariance sizing. The dominant force in the LEO orbit regime is the drag force caused by atmospheric drag. Accurate accounting of the drag force is especially important for the accuracy of the propagated state. The implementation of a box and wing model to improve drag estimation accuracy, and its overall effect on the covariance state is explored.The process of tuning the EKF for Aqua and Aura support is described, including examination of the measurement errors of available observation types (Doppler and range), and methods of dealing with potentially volatile atmospheric drag modeling. Predictive accuracy and the distribution of the Chi-square statistic, calculated based of the ODTK EKF solutions, are assessed versus accepted norms for the orbit regime.

Quasi-Tangency Points on the Orbits of a Small Body and a Planet at the Low-Velocity Encounter

NASA Astrophysics Data System (ADS)

Emel'yanenko, N. Yu.

2018-03-01

We propose a method for selecting a low-velocity encounter of a small body with a planet from the evolution of the orbital elements. Polar orbital coordinates of the quasi-tangency point on the orbit of a small body are determined. Rectangular heliocentric coordinates of the quasi-tangency point on the orbit of a planet are determined. An algorithm to search for low-velocity encounters in the evolution of the orbital elements of small bodies is described. The low-velocity encounter of comet 39P/Oterma with Jupiter is considered as an example.

Station coordinates, baselines, and earth rotation from Lageos laser ranging - 1976-1984

NASA Technical Reports Server (NTRS)

Tapley, B. D.; Schultz, B. E.; Eanes, R. J.

1985-01-01

The orbit of the Lageos satellite is well suited as a reference frame for studying the rotation of the earth and the relative motion of points on the earth's crust. The satellite laser measurements can determine the location of a set of tracking stations in an appropriate terrestrial coordinate system. The motion of the earth's rotation axis relative to this system can be studied on the basis of the established tracking station locations. The present investigation is concerned with an analysis of 7.7 years of Lageos laser ranging data. In the first solution considered, the entire data span was used to adjust a single set of station positions simultaneously with orbit and earth rotation parameters. Attention is given to the accuracy of earth rotation parameters which are determined as an inherent part of the solution process.

A simple criterion for determining the dynamical stability of three-body systems

NASA Technical Reports Server (NTRS)

Black, D. C.

1982-01-01

Coplanar, prograde three-body systems (TBS) are discussed, emphasizing the specification of general criteria for determining whether such systems are dynamically stable. It is shown that the Graziani-Black (1981) criteria provide a quantitatively accurate characterization of the onset of dynamic instability for values of the dimensionless mass ranging from one millionth to one million. Harrington's (1977) general criterion and the Graziani-Black criterion are compared with results from analytic work that spans a 12-orders-of-magnitude variation in the mass ratios of the TBS components. Comparison of the Graziani-Black criteria with data for eight well-studied triple-star systems indicates that the observed lower limit for the ratio of periastron distance of the tertiary orbit to the semimajor axis of the binary orbit is due to dynamical instability rather than to cosmogonic processes.

Seasat. Volume 4: Attitude determination

NASA Technical Reports Server (NTRS)

Treder, A. J.

1980-01-01

The Seasat project was a feasibility demonstration of the use of orbital remote sensing for global ocean observation. The satellite was launched in June 1978 and was operated successfully until October 1978. A massive electrical failure occurred in the power system, terminating the mission prematurely. The actual implementation of the Seasat Attitude Determination system and the contents of the attitude data files generated by that system are documented. The deviations from plan caused by the anomalous Sun interference with horizon sensors, inflight calibration of Sun sensor head 2 alignment and horizon sensor biomass, estimation of yaw interpolation parameters, Sun and horizon sensor error sources, and yaw interpolation accuracy are included. Examples are given of flight attitude data from all modes of the Orbital Attitude Control System, of the ground processing effects on attitude data, and of cold cloud effects on pitch, and roll data.

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

NASA Technical Reports Server (NTRS)

Auziasdeturenne, Jacqueline; Beall, Mark; Burianek, Joseph; Cinniger, Anna; Dunmire, Barbrina; Haberman, Eric; Iwamoto, James; Johnson, Stephen; Mccracken, Shawn; Miller, Melanie

1989-01-01

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Baryonic impact on the dark matter orbital properties of Milky Way-sized haloes

NASA Astrophysics Data System (ADS)

Zhu, Qirong; Hernquist, Lars; Marinacci, Federico; Springel, Volker; Li, Yuexing

2017-04-01

We study the orbital properties of dark matter haloes by combining a spectral method and cosmological simulations of Milky Way-sized Galaxies. We compare the dynamics and orbits of individual dark matter particles from both hydrodynamic and N-body simulations, and find that the fraction of box, tube and resonant orbits of the dark matter halo decreases significantly due to the effects of baryons. In particular, the central region of the dark matter halo in the hydrodynamic simulation is dominated by regular, short-axis tube orbits, in contrast to the chaotic, box and thin orbits dominant in the N-body run. This leads to a more spherical dark matter halo in the hydrodynamic run compared to a prolate one as commonly seen in the N-body simulations. Furthermore, by using a kernel-based density estimator, we compare the coarse-grained phase-space densities of dark matter haloes in both simulations and find that it is lower by ˜0.5 dex in the hydrodynamic run due to changes in the angular momentum distribution, which indicates that the baryonic process that affects the dark matter is irreversible. Our results imply that baryons play an important role in determining the shape, kinematics and phase-space density of dark matter haloes in galaxies.

PCVs Estimation and their Impacts on Precise Orbit Determination of LEOs

NASA Astrophysics Data System (ADS)

Chunmei, Z.; WANG, X.

2017-12-01

In the last decade the precise orbit determination (POD) based on GNSS, such as GPS, has been considered as one of the efficient methods to derive orbits of Low Earth Orbiters (LEOs) that demand accuracy requirements. The Earth gravity field recovery and its related researches require precise dynamic orbits of LEOs. With the improvements of GNSS satellites' orbit and clock accuracy, the algorithm optimization and the refinement of perturbation force models, the antenna phase-center variations (PCVs) of space-borne GNSS receiver have become an increasingly important factor that affects POD accuracy. A series of LEOs such as HY-2, ZY-3 and FY-3 with homebred space-borne GNSS receivers have been launched in the past several years in China. Some of these LEOs load dual-mode GNSS receivers of GPS and BDS signals. The reliable performance of these space-borne receivers has been establishing an important foundation for the future launches of China gravity satellites. Therefore, we first evaluate the data quality of on-board GNSS measurement by examining integrity, multipath error, cycle slip ratio and other quality indices. Then we determine the orbits of several LEOs at different altitudes by the reduced dynamic orbit determination method. The corresponding ionosphere-free carrier phase post-fit residual time series are obtained. And then we establish the PCVs model by the ionosphere-free residual approach and analyze the effects of antenna phase-center variation on orbits. It is shown that orbit accuracy of LEO satellites is greatly improved after in-flight PCV calibration. Finally, focus on the dual-mode receiver of FY-3 satellite we analyze the quality of onboard BDS data and then evaluate the accuracy of the FY-3 orbit determined using only BDS measurement onboard. The accuracy of LEO satellites orbit based on BDS would be well improved with the global completion of BDS by 2020.

Generating precise and homogeneous orbits for Jason-1 and Jason-2

NASA Astrophysics Data System (ADS)

Flohrer, Claudia; Otten, Michiel; Springer, Tim; Dow, John

2011-07-01

Driven by the GMES (Global Monitoring for Environment and Security) and GGOS (Global Geodetic Observing System) initiatives the user community has a strong demand for high-quality altimetry products. In order to derive such high-quality altimetry products, precise orbits for the altimetry satellites are a necessity. With the launch of the TOPEX/Poseidon mission in 1992 a still on-going time series of high-accuracy altimetry measurements of ocean topography started, continued by the altimetry missions Jason-1 in 2001 and Jason-2/OSTM in 2008. This paper contributes to the on-going orbit reprocessing carried out by several groups and presents the efforts of the Navigation Support Office at ESA/ESOC using its NAPEOS software for the generation of precise and homogeneous orbits referring to the same reference frame for the altimetry satellites Jason-1 and Jason-2. Data of all three tracking instruments on-board the satellites (beside the altimeter), i.e. GPS, DORIS, and SLR measurements, were used in a combined data analysis. About 7 years of Jason-1 data and more than 1 year of Jason-2 data were processed. Our processing strategy is close to the GDR-C standards. However, we estimated slightly different scaling factors for the solar radiation pressure model of 0.96 and 0.98 for Jason-1 and Jason-2, respectively. We used 30 s sampled GPS data and introduced 30 s satellite clocks stemming from ESOC's reprocessing of the combined GPS/GLONASS IGS solution. We present the orbit determination results, focusing on the benefits of adding GPS data to the solution. The fully combined solution was found to give the best orbit results. We reach a post-fit RMS of the GPS phase observation residuals of 6 mm for Jason-1 and 7 mm for Jason-2. The DORIS post-fit residuals clearly benefit from using GPS data in addition, as the DORIS data editing improves. The DORIS observation RMS for the fully combined solution is with 3.5 mm and 3.4 mm, respectively, 0.3 mm better than for the DORIS-SLR solution. Our orbit solution agrees well with external solutions from other analysis centers, as CNES, LCA, and JPL. The orbit differences between our fully combined orbits and the CNES GDR-C orbits are of about 0.8 cm for Jason-1 and at 0.9 cm for Jason-2 in the radial direction. In the cross-track component we observe a clear improvement when adding GPS data to the POD process. The 3D-RMS of the orbit differences reveals a good orbit consistency at 2.7 cm and 2.9 cm for Jason-1 and Jason-2. Our resulting orbit series for both Jason satellites refer to the ITRF2005 reference frame and are provided in sp3 file format on our ftp server.

Determination of celestial bodies orbits and probabilities of their collisions with the Earth

NASA Astrophysics Data System (ADS)

Medvedev, Yuri; Vavilov, Dmitrii

In this work we have developed a universal method to determine the small bodies orbits in the Solar System. In the method we consider different planes of body’s motion and pick up which is the most appropriate. Given an orbit plane we can calculate geocentric distances at time of observations and consequence determinate all orbital elements. Another technique that we propose here addresses the problem of estimation probability of collisions celestial bodies with the Earth. This technique uses the coordinate system associated with the nominal osculating orbit. We have compared proposed technique with the Monte-Carlo simulation. Results of these methods exhibit satisfactory agreement, whereas, proposed method is advantageous in time performance.

STS-28 Columbia, Orbiter Vehicle (OV) 102, roll over to KSC OPF

NASA Image and Video Library

1989-01-24

S89-28090 (24 Jan 1989) --- The orbiter Columbia, slated for mission STS-28 is rolled over from the Orbiter Maintenance and Refurbishment Facility to the Orbiter Processing Facility bay 1 to begin processing for its mission in July 1989. The Department of Defense dedicted mission will be commanded by astronaut Brewster Shaw.

General view from outside the Orbiter Processing Facility at the ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view from outside the Orbiter Processing Facility at the Kennedy Space Center with the bay doors open as the Orbiter Discovery is atop the transport vehicle prepared to be moved over to the Vehicle Assembly Building. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, Richard Parker, with NASA, watches a monitor showing images from a camera inserted beneath tiles of the orbiter Endeavour to inspect for corrosion.

NASA Image and Video Library

2003-09-04

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, Richard Parker, with NASA, watches a monitor showing images from a camera inserted beneath tiles of the orbiter Endeavour to inspect for corrosion.

Satellite laser ranging to low Earth orbiters: orbit and network validation

NASA Astrophysics Data System (ADS)

Arnold, Daniel; Montenbruck, Oliver; Hackel, Stefan; Sośnica, Krzysztof

2018-04-01

Satellite laser ranging (SLR) to low Earth orbiters (LEOs) provides optical distance measurements with mm-to-cm-level precision. SLR residuals, i.e., differences between measured and modeled ranges, serve as a common figure of merit for the quality assessment of orbits derived by radiometric tracking techniques. We discuss relevant processing standards for the modeling of SLR observations and highlight the importance of line-of-sight-dependent range corrections for the various types of laser retroreflector arrays. A 1-3 cm consistency of SLR observations and GPS-based precise orbits is demonstrated for a wide range of past and present LEO missions supported by the International Laser Ranging Service (ILRS). A parameter estimation approach is presented to investigate systematic orbit errors and it is shown that SLR validation of LEO satellites is not only able to detect radial but also along-track and cross-track offsets. SLR residual statistics clearly depend on the employed precise orbit determination technique (kinematic vs. reduced-dynamic, float vs. fixed ambiguities) but also reveal pronounced differences in the ILRS station performance. Using the residual-based parameter estimation approach, corrections to ILRS station coordinates, range biases, and timing offsets are derived. As a result, root-mean-square residuals of 5-10 mm have been achieved over a 1-year data arc in 2016 using observations from a subset of high-performance stations and ambiguity-fixed orbits of four LEO missions. As a final contribution, we demonstrate that SLR can not only validate single-satellite orbit solutions but also precise baseline solutions of formation flying missions such as GRACE, TanDEM-X, and Swarm.

The origin of phosphorescence in Iridium (III) complexes. The role of relativistic effects

NASA Astrophysics Data System (ADS)

Cantero-López, Plinio; Páez-Hernández, Dayan; Arratia-Pérez, Ramiro

2017-10-01

A series of luminescent Ir(III) complexes of the type [Ir(F2ppy)2L] (where L = Lpytz , LOMe , Lbut) have been studied using relativistic two-component density functional theory considering the spin-orbit coupling. The absorption spectra of the three complexes were determined. The most important transition appears in the region between 250 and 350 nm, which is in good agreement with the experimental reports. The three complexes show phosphorescent properties due to a metal-ligand charge transfer (MLCT) process, where the spin-orbit coupling (SOC) plays a key role due to the introduction of a zero field splitting (ZFS) and the mixing of states with different spins which contributes to modify the emission selection rule. The lifetimes of the emission processes were calculated, and the values are in the same order of the experimental reports.

Passive correlation ranging of a geostationary satellite using DVB-S payload signals.

NASA Astrophysics Data System (ADS)

Shakun, Leonid; Shulga, Alexandr; Sybiryakova, Yevgeniya; Bushuev, Felix; Kaliuzhnyi, Mykola; Bezrukovs, Vladislavs; Moskalenko, Sergiy; Kulishenko, Vladislav; Balagura, Oleg

2016-07-01

Passive correlation ranging (PaCoRa) for geostationary satellites is now considered as an alternate to tone-ranging (https://artes.esa.int/search/node/PaCoRa). The PaCoRa method has been employed in the Research Institute "Nikolaev astronomical observatory" since the first experiment in August 2011 with two stations spatially separated on 150 km. The PaCoRa has been considered as an independent method for tracking the future Ukrainian geostationary satellite "Lybid'. Now a radio engineering complex (RC) for passive ranging consists of five spatially separated stations of receiving digital satellite television and a data processing center located in Mykolaiv. The stations are located in Kyiv, Kharkiv, Mukacheve, Mykolaiv (Ukraine) and in Ventspils (Latvia). Each station has identical equipment. The equipment allows making synchronous recording of fragments of the DVB-S signal from the quadrature detector output of a satellite television receiver. The fragments are recorded every second. Synchronization of the stations is performed using GPS receivers. Samples of the complex signal obtained in this way are archived and are sent to the data processing center over the Internet. Here the time differences of arrival (TDOA) for pairs of the stations are determined as a result of correlation processing of received signals. The values of the TDOA that measured every second are used for orbit determination (OD) of the satellite. The results of orbit determination of the geostationary telecommunication satellite "Eutelsat-13B" (13º East) obtained during about four months of observations in 2015 are presented in the report. The TDOA and OD accuracies are also given. Single-measurement error (1 sigma) of the TDOA is equal about 8.7 ns for all pairs of the stations. Standard deviations and average values of the residuals between the observed TDOA and the TDOA computed using the orbit elements obtained from optical measurements are estimated for the pairs Kharkiv-Mykolaiv and Mukacheve-Mykolaiv. The standard deviations do not exceed 10 ns for the both pairs and the average values are equal +10 ns and -106 ns respectively for Kharkiv-Mykolaiv and Mukacheve-Mykolaiv. We discuss the residuals between the observed TDOA and estimates of the TDOA that are calculated by fitted models of satellite motion: the SGP4/SDP4 model and the model based on the numerical integration of the equations of motion taking into account the geopotential, and the perturbation from the Moon and the Sun. We note that residuals from the model SGP4/SDP4 have periodic deviations due to the inaccuracy of the SGP4/SDP4 model. As a result, estimation of the standard deviation of the satellite position is about 60 m for the epoch of the SGP4/SDP4 orbit elements. The residuals for the numerical model in the interval of one day do not show low-frequency deviation. In this case, the estimate of the standard deviation of the satellite position is about 12 m for the epoch of the numerical orbit elements. Keywords. DVB-S, geostationary satellite, orbit determination, passive ranging.

Orbit Determination (OD) Error Analysis Results for the Triana Sun-Earth L1 Libration Point Mission and for the Fourier Kelvin Stellar Interferometer (FKSI) Sun-Earth L2 Libration Point Mission Concept

NASA Technical Reports Server (NTRS)

Marr, Greg C.

2003-01-01

The Triana spacecraft was designed to be launched by the Space Shuttle. The nominal Triana mission orbit will be a Sun-Earth L1 libration point orbit. Using the NASA Goddard Space Flight Center's Orbit Determination Error Analysis System (ODEAS), orbit determination (OD) error analysis results are presented for all phases of the Triana mission from the first correction maneuver through approximately launch plus 6 months. Results are also presented for the science data collection phase of the Fourier Kelvin Stellar Interferometer Sun-Earth L2 libration point mission concept with momentum unloading thrust perturbations during the tracking arc. The Triana analysis includes extensive analysis of an initial short arc orbit determination solution and results using both Deep Space Network (DSN) and commercial Universal Space Network (USN) statistics. These results could be utilized in support of future Sun-Earth libration point missions.

On-Orbit ACDS Performance of the Landsat 7 Spacecraft

NASA Technical Reports Server (NTRS)

Sabelhaus, Phillip; Bolek, Joseph; Scott, Steve; Holmes, Eric; O'Donnell, James R., Jr.; Storey, James

2001-01-01

Landsat 7 is part of NASA's Earth Science Enterprise (ESE). The ESE is committed to developing an understanding of the total Earth system, the effects of natural and human-induced changes on the global environment, and how natural processes affect humans and how humans affect them. The Landsat 7 satellite consists of the spacecraft bus which was provided under a NASA contract with Lockheed Martin Missiles and Space in Philadelphia, PA, and the Enhanced Thematic Mapper-Plus (ETM+) instrument, procured under a NASA contract with Raytheon Santa Barbara Remote Sensing, in Santa Barbara, CA. The Landsat 7 Attitude Control and Determination System (ACDS) provides many essential functions for the operation of the spacecraft bus and for ETM+. The ACDS maintains the required attitude and orbit at the degree of accuracy necessary for power generation, command and telemetry, thermal balance, image acquisition, Gimbaled X-Band Antenna (GXA) pointing and data for image post-processing. Descriptions of the Landsat 7 mission and the ACDS modes and requirements are presented. A brief summary of significant events of the on-orbit initialization and validation period are provided. Finally, the Landsat 7 product generation system is described and the impact that the ACDS performance has on the ground based image processing system is explored.

Using an Iterative Fourier Series Approach in Determining Orbital Elements of Detached Visual Binary Stars

NASA Astrophysics Data System (ADS)

Tupa, Peter R.; Quirin, S.; DeLeo, G. G.; McCluskey, G. E., Jr.

2007-12-01

We present a modified Fourier transform approach to determine the orbital parameters of detached visual binary stars. Originally inspired by Monet (ApJ 234, 275, 1979), this new method utilizes an iterative routine of refining higher order Fourier terms in a manner consistent with Keplerian motion. In most cases, this approach is not sensitive to the starting orbital parameters in the iterative loop. In many cases we have determined orbital elements even with small fragments of orbits and noisy data, although some systems show computational instabilities. The algorithm was constructed using the MAPLE mathematical software code and tested on artificially created orbits and many real binary systems, including Gliese 22 AC, Tau 51, and BU 738. This work was supported at Lehigh University by NSF-REU grant PHY-9820301.

Topography of the Lunar Poles and Application to Geodesy with the Lunar Reconnaissance Orbiter

NASA Technical Reports Server (NTRS)

Mazarico, Erwan; Neumann, Gregory A.; Rowlands, David D.; Smith, David E.; Zuber, Maria T.

2012-01-01

The Lunar Orbiter Laser Altimeter (LOLA) [1] onboard the Lunar Reconnaissance Orbiter (LRO) [2] has been operating continuously since July 2009 [3], accumulating approx.5.4 billion measurements from 2 billion on-orbit laser shots. LRO s near-polar orbit results in very high data density in the immediate vicinity of the lunar poles, which are each sampled every 2h. With more than 10,000 orbits, high-resolution maps can be constructed [4] and studied [5]. However, this requires careful processing of the raw data, as subtle errors in the spacecraft position and pointing can lead to visible artifacts in the final map. In other locations on the Moon, ground tracks are subparallel and longitudinal separations are typically a few hundred meters. Near the poles, the track intersection angles can be large and the inter-track spacing is small (above 80 latitude, the effective resolution is better than 50m). Precision Orbit Determination (POD) of the LRO spacecraft [6] was performed to satisfy the LOLA and LRO mission requirements, which lead to a significant improvement in the orbit position knowledge over the short-release navigation products. However, with pixel resolutions of 10 to 25 meters, artifacts due to orbit reconstruction still exist. Here, we show how the complete LOLA dataset at both poles can be adjusted geometrically to produce a high-accuracy, high-resolution maps with minimal track artifacts. We also describe how those maps can then feedback to the POD work, by providing topographic base maps with which individual LOLA altimetric measurements can be contributing to orbit changes. These direct altimetry constraints improve accuracy and can be used more simply than the altimetric crossovers [6].

Determination of Orbiter and Carrier Aerodynamic Coefficients from Load Cell Measurements

NASA Technical Reports Server (NTRS)

Glenn, G. M.

1976-01-01

A method of determining orbiter and carrier total aerodynamic coefficients from load cell measurements is required to support the inert and the captive active flights of the ALT program. A set of equations expressing the orbiter and carrier total aerodynamic coefficients in terms of the load cell measurements, the sensed dynamics of the Boeing 747 (carrier) aircraft, and the relative geometry of the orbiter/carrier is derived.

General View looking forward along the centerline of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General View looking forward along the centerline of the Orbiter Discovery looking into the payload bay with a payload in the process of being secured into place. This photograph was taken in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Orbit-product representation and correction of Gaussian belief propagation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Johnson, Jason K; Chertkov, Michael; Chernyak, Vladimir

We present a new interpretation of Gaussian belief propagation (GaBP) based on the 'zeta function' representation of the determinant as a product over orbits of a graph. We show that GaBP captures back-tracking orbits of the graph and consider how to correct this estimate by accounting for non-backtracking orbits. We show that the product over non-backtracking orbits may be interpreted as the determinant of the non-backtracking adjacency matrix of the graph with edge weights based on the solution of GaBP. An efficient method is proposed to compute a truncated correction factor including all non-backtracking orbits up to a specified length.

MERCATOR: Methods and Realization for Control of the Attitude and the Orbit of spacecraft

NASA Technical Reports Server (NTRS)

Tavernier, Gilles; Campan, Genevieve

1993-01-01

Since 1974, CNES has been involved in geostationary positioning. Among different entities participating in operations and their preparation, the Flight Dynamics Center (FDC) is in charge of performing the following tasks: orbit determination; attitude determination; computation, monitoring, and calibration of orbit maneuvers; computation, monitoring, and calibration of attitude maneuvers; and operational predictions. In order to fulfill this mission, the FDC receives telemetry from the satellite and localization measurements from ground stations (e.g., CNES, NASA, INTELSAT). These data are processed by space dynamics programs integrated in the MERCATOR system which is run on SUN workstations (UNIX O.S.). The main features of MERCATOR are redundancy, modularity, and flexibility: efficient, flexible, and user friendly man-machine interface; and four identical SUN stations redundantly linked in an Ethernet network. Each workstation can perform all the tasks from data acquisition to computation results dissemination through a video network. A team of four engineers can handle the space mechanics aspects of a complete geostationary positioning from the injection into a transfer orbit to the final maneuvers in the station-keeping window. MERCATOR has been or is to be used for operations related to more than ten geostationary positionings. Initially developed for geostationary satellites, MERCATOR's methodology was also used for satellite control centers and can be applied to a wide range of satellites and to future manned missions.

Independent Orbiter Assessment (IOA): Analysis of the body flap subsystem

NASA Technical Reports Server (NTRS)

Wilson, R. E.; Riccio, J. R.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items (PCIs). To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Orbiter Body Flap (BF) subsystem hardware are documented. The BF is a large aerosurface located at the trailing edge of the lower aft fuselage of the Orbiter. The proper function of the BF is essential during the dynamic flight phases of ascent and entry. During the ascent phase of flight, the BF trails in a fixed position. For entry, the BF provides elevon load relief, trim control, and acts as a heat shield for the main engines. Specifically, the BF hardware comprises the following components: Power Drive Unit (PDU), rotary actuators, and torque tubes. The IOA analysis process utilized available BF hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 35 failure modes analyzed, 19 were determined to be PCIs.

An optical survey for space debris on highly eccentric and inclined MEO orbits

NASA Astrophysics Data System (ADS)

Schildknecht, Thomas; Flohrer, Tim; Hinze, Andreas; Vananti, Alessandro; Silha, Jiri

Optical surveys for space debris in high-altitude orbits have been conducted since more than fifteen years. Originally these efforts concentrated mainly on the geostationary ring (GEO) and its close region. Corresponding observation strategies, processing techniques and cataloguing approaches have been developed and successfully applied. The ESA GEO surveys, e.g., resulted in the detection of a significant population of small-size debris and later in the discovery of high area-to-mass ratio objects in GEO-like orbits. The observation scenarios were successively adapted to survey the geostationary transfer orbit (GTO) region; and surveys to search for debris in the medium Earth orbit (MEO) region of the global navigation satellite constellations were successfully conducted. Comparably less experience (both, in terms of practical observation and strategy definition) is available for eccentric orbits that (at least partly) are in the MEO region, in particular for the Molniya-type orbits. Several breakup events and deliberate fragmentations are known to have taken place in such orbits. Survey and follow-up strategies for searching space debris objects in highly-eccentric MEO orbits, and to acquire orbits which are sufficiently accurate to catalogue such objects and to maintain their orbits over longer time spans were developed and, eventually, optical observations were conducted in the framework of an ESA study using ESA' Space Debris Telescope (ESASDT) the 1-m Zeiss telescope located at the Optical Ground Station (OGS) at the Teide Observatory at Tenerife, Spain. Thirteen nights of surveys of Molniya-type orbits was performed between January and August 2013. A basic survey consisted of observing a single geocentric field for 10 minutes. If a faint object was found, follow-up observations were performed during the same night to ensure a save rediscovery of the object during the next nights. Additional follow-up observations to maintain the orbits of these newly discovered faint objects were also acquired with AIUB's 1 m ZIMLAT telescope in Zimmerwald, Switzerland. Eventually, 255 basic surveys were performed during these thirteen nights corresponding to about 47 hours of observations. In total 30 uncorrelated faint objects were discovered. On average one uncorrelated object was found every 100 minutes. Some of these objects show a considerable brightness variation and have a high area-to-mass ratio as determined in the orbit estimation process. We also investigated the detection efficiency of our surveys by comparing the observation results with the TLE population by means of ESA's PROOF tool. Furthermore a comparison of the real detections with the statistical population of the ESA MASTER-2009 model was performed. The result shows that the fragment population of objects in Molniya-type orbits is underestimated in the MASTER model.

The Ames-Lockheed orbiter processing scheduling system

NASA Technical Reports Server (NTRS)

Zweben, Monte; Gargan, Robert

1991-01-01

A general purpose scheduling system and its application to Space Shuttle Orbiter Processing at the Kennedy Space Center (KSC) are described. Orbiter processing entails all the inspection, testing, repair, and maintenance necessary to prepare the Shuttle for launch and takes place within the Orbiter Processing Facility (OPF) at KSC, the Vehicle Assembly Building (VAB), and on the launch pad. The problems are extremely combinatoric in that there are thousands of tasks, resources, and other temporal considerations that must be coordinated. Researchers are building a scheduling tool that they hope will be an integral part of automating the planning and scheduling process at KSC. The scheduling engine is domain independent and is also being applied to Space Shuttle cargo processing problems as well as wind tunnel scheduling problems.

System design and instrument development for future formation-flying magnetospheric satellite mission SCOPE

NASA Astrophysics Data System (ADS)

Saito, Y.; Fujimoto, M.; Maezawa, K.; Kojima, H.; Takashima, T.; Matsuoka, A.; Shinohara, I.; Tsuda, Y.; Higuchi, K.; Toda, T.

Japan Aerospace Exploration Agency JAXA is currently planning a next generation magnetosphere observation mission called SCOPE cross-Scale COupling in the Plasma universE The main purpose of this mission is to investigate the dynamic behaviors of plasmas in the Terrestrial magnetosphere that range over various time and spatial scales The basic idea of the SCOPE mission is to distinguish temporal and spatial variations of physical processes by putting five formation flying spacecraft into the key region of the Terrestrial magnetosphere The orbit of SCOPE is a highly elliptical orbit with its apogee 30Re from the Earth center SCOPE consists of one 450kg mother satellite and four 90kg daughter satellites flying 5 to 5000km apart from each other The inter-satellite link is used for telemetry command operation as well as ranging to determine the relative orbit of 5 satellites in a small distance which cannot be resolved by the ground-based orbit determination The SCOPE mission is designed such that observational studies from the new perspective that is the cross-scale coupling viewpoint are enabled The orbit is so designed that the spacecraft will visit most of the key regions in the magnetosphere that is the bow shock the magnetospheric boundary the inner-magnetosphere and the near-Earth magnetotail In order to realize the science objectives high performance Plasma Particle sensors DC AC Magnetic and Electric field sensors and Wave Particle Correlator are planned to be onboard the SCOPE satellite All the SCOPE satellites have two 5m spin-axis antenna

Development of an integrated spacecraft Guidance, Navigation, & Control subsystem for automated proximity operations

NASA Astrophysics Data System (ADS)

Schulte, Peter Z.; Spencer, David A.

2016-01-01

This paper describes the development and validation process of a highly automated Guidance, Navigation, & Control subsystem for a small satellite on-orbit inspection application, enabling proximity operations without human-in-the-loop interaction. The paper focuses on the integration and testing of Guidance, Navigation, & Control software and the development of decision logic to address the question of how such a system can be effectively implemented for full automation. This process is unique because a multitude of operational scenarios must be considered and a set of complex interactions between subsystem algorithms must be defined to achieve the automation goal. The Prox-1 mission is currently under development within the Space Systems Design Laboratory at the Georgia Institute of Technology. The mission involves the characterization of new small satellite component technologies, deployment of the LightSail 3U CubeSat, entering into a trailing orbit relative to LightSail using ground-in-the-loop commands, and demonstration of automated proximity operations through formation flight and natural motion circumnavigation maneuvers. Operations such as these may be utilized for many scenarios including on-orbit inspection, refueling, repair, construction, reconnaissance, docking, and debris mitigation activities. Prox-1 uses onboard sensors and imaging instruments to perform Guidance, Navigation, & Control operations during on-orbit inspection of LightSail. Navigation filters perform relative orbit determination based on images of the target spacecraft, and guidance algorithms conduct automated maneuver planning. A slew and tracking controller sends attitude actuation commands to a set of control moment gyroscopes, and other controllers manage desaturation, detumble, thruster firing, and target acquisition/recovery. All Guidance, Navigation, & Control algorithms are developed in a MATLAB/Simulink six degree-of-freedom simulation environment and are integrated using decision logic to autonomously determine when actions should be performed. The complexity of this decision logic is the primary challenge of the automated process, and the Stateflow tool in Simulink is used to establish logical relationships and manage data flow between each of the individual hardware and software components. Once the integrated simulation is fully developed in MATLAB/Simulink, the algorithms are autocoded to C/C++ and integrated into flight software. Hardware-in-the-loop testing provides validation of the Guidance, Navigation, & Control subsystem performance.

Computer program determines thermal environment and temperature history of lunar orbiting space vehicles

NASA Technical Reports Server (NTRS)

Head, D. E.; Mitchell, K. L.

1967-01-01

Program computes the thermal environment of a spacecraft in a lunar orbit. The quantities determined include the incident flux /solar and lunar emitted radiation/, total radiation absorbed by a surface, and the resulting surface temperature as a function of time and orbital position.

Closeup view of the aft fuselage of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up view of the aft fuselage of the Orbiter Discovery on the starboard side looking forward. This view is of the attach surface for the Orbiter Maneuvering System/Reaction Control System (OMS/RCS) Pod. The OMS/RCS pods are removed for processing and reconditioning at another facility. This view was taken from a service platform in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Shuttle OFT Level C navigation requirements

NASA Technical Reports Server (NTRS)

1980-01-01

Detailed requirements for the orbital operations computer loads, OPS 2, and OPS 8 are given. These requirements represent the total on-orbit/rendezvous navigation baseline requirements for the following principal functions: on-orbital/rendezvous navigation sequencer; on-orbit/rendezvous UPP sequencer; on-orbit rendezvous navigation; on-orbit prediction; on-orbit user parameter processing; and landing Site update.

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility insert the liquid oxygen feedline for the 17-inch disconnect in the orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility insert the liquid oxygen feedline for the 17-inch disconnect in the orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers install the liquid oxygen feedline for the 17-inch disconnect on orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers install the liquid oxygen feedline for the 17-inch disconnect on orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers raise the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers raise the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers lift the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers lift the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers move the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers move the liquid oxygen feedline for the 17-inch disconnect toward orbiter Discovery for installation. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

Geosynchronous Orbit Determination Using Space Surveillance Network Observations and Improved Radiative Force Modeling

DTIC Science & Technology

2004-06-01

equinoctial elements , because both sets of orbital elements reference the equinoctial coordinate system. In fact, to...spacecraft position and velocity vectors, or an element set , which represents the orbit using scalar quantities and angle measurements called orbital ...common element sets used to describe elliptical orbits (including circular orbits ) are Keplerian elements , also called classical orbital

Planning Bepicolombo MPO Science Operations to study Mercury Interior

NASA Astrophysics Data System (ADS)

De La Fuente, Sara; Carasa, Angela; Ortiz, Iñaki; Rodriguez, Pedro; Casale, Mauro; Benkhoff, Johannes; Zender, Joe

2017-04-01

BepiColombo is an Interdisciplinary Cornerstone ESA-JAXA Mission to Mercury, with two orbiters, the ESA Mercury Planetary Orbiter (MPO) and the JAXA Mercury Magnetospheric Orbiter (MMO) dedicated to study of the planet and its magnetosphere. The MPO, is a three-axis-stabilized, nadir-pointing spacecraft which will be placed in a polar orbit, providing excellent spatial resolution over the entire planet surface. The MPO's scientific payload comprises 11 instrument packages, including laser altimeter, cameras and the radio science experiment that will be dedicated to the study of Mercury's interior: structure, composition, formation and evolution. The planning of the science operations to be carried out by the Mercury's interior scientific instruments will be done by the SGS located at the European Space Astronomy Centre (ESAC), in conjunction with the scientific instrument teams. The process will always consider the complete nominal mission duration, such that the contribution of the scheduled science operations to the science objectives, the total data volume generated, and the seasonal interdependency, can be tracked. The heart of the science operations planning process is the Observations Catalogue (OC), a web-accessed database to collect and analyse all science operations requests. From the OC, the SGS will first determine all science opportunity windows compatible with the spacecraft operational constraints. Secondly, only those compatible with the resources (power and data volume) and pointing constraints will be chosen, including slew feasibility.

Implementing a 50x50 Gravity Field Model in an Orbit Determination System

DTIC Science & Technology

1993-06-01

orbital element set , sometimes better known as the Keplerian orbital element set . Another set is the equinoctial element set , which removes singularity...Conference. San Diego, California. August 1976. [8] Cefola, Paul. Equinoctial Orbit Elements - Application to Artificial Satellite Orbits . AIAA Paper...251 A.2 Classical Orbital Elements ......................................................... 251 A.3

High Degree and Order Gravity Fields of the Moon Derived from GRAIL Data

NASA Technical Reports Server (NTRS)

Lemoine, F. G.; Goossens, S. J.; Sabaka, T. J.; Nicholas, J. B.; Mazarico, E.; Rowlands, D. D.; Loomis, B. D.; Chinn, D. S.; Caprette, D. S.; McCarthy, J. J.;

Lunar base mission technology issues and orbital demonstration requirements on space station

NASA Technical Reports Server (NTRS)

Llewellyn, Charles P.; Weidman, Deene J.

1992-01-01

The International Space Station has been the object of considerable design, redesign, and alteration since it was originally proposed in early 1984. In the intervening years the station has slowly evolved to a specific design that was thoroughly reviewed by a large agency-wide Critical Evaluation Task Force (CETF). As space station designs continue to evolve, studies must be conducted to determine the suitability of the current design for some of the primary purposes for which the station will be used. This paper concentrates on the technology requirements and issues, the on-orbit demonstration and verification program, and the space station focused support required prior to the establishment of a permanently manned lunar base as identified in the National Commission on Space report. Technology issues associated with the on-orbit assembly and processing of the lunar vehicle flight elements are also discussed.

Magnetization reversal in ferromagnetic thin films induced by spin-orbit interaction with Slonczewski-like spin transfer torque

DOE Office of Scientific and Technical Information (OSTI.GOV)

Li, Jia, E-mail: lijia@wipm.ac.cn

2014-10-07

We theoretically investigate the dynamics of magnetization in ferromagnetic thin films induced by spin-orbit interaction with Slonczewski-like spin transfer torque. We reproduce the experimental results of perpendicular magnetic anisotropy films by micromagnetic simulation. Due to the spin-orbit interaction, the magnetization can be switched by changing the direction of the current with the assistant of magnetic field. By increasing the current amplitude, wider range of switching events can be achieved. Time evolution of magnetization has provided us a clear view of the process, and explained the role of minimum external field. Slonczewski-like spin transfer torque modifies the magnetization when current ismore » present. The magnitude of the minimum external field is determined by the strength of the Slonczewski-like spin transfer torque. The investigations may provide potential applications in magnetic memories.« less

Navigation of space VLBI missions: Radioastron and VSOP

NASA Technical Reports Server (NTRS)

Ellis, Jordan

1993-01-01

In the mid-1990s, Russian and Japanese space agencies will each place into highly elliptic earth orbit a radio telescope consisting of a large antenna and radio astronomy receivers. Very long baseline interferometry (VLBI) techniques will be used to obtain high resolution images of radio sources observed by the space and ground based antennas. Stringent navigation accuracy requirements are imposed on the space VLBI missions by the need to transfer an ultra-stable ground reference frequency standard to the spacecraft and by the demands of the VLBI correlation process. Orbit determination for the mission will be the joint responsibility of navigation centers in the U.S., Russia, and Japan with orbit estimates based on combining tracking data from NASA, Russian, and Japanese sites. This paper describes the operational plans, the inter-agency coordination, and data exchange between the navigation centers required for space VLBI navigation.

Contamination control of the space shuttle Orbiter crew compartment

NASA Technical Reports Server (NTRS)

Bartelson, Donald W.

1986-01-01

Effective contamination control as applied to manned space flight environments is a discipline characterized and controlled by many parameters. An introduction is given to issues involving Orbiter crew compartment contamination control. An effective ground processing contamination control program is an essential building block to a successful shuttle mission. Personnel are required to don cleanroom-grade clothing ensembles before entering the crew compartment and follow cleanroom rules and regulations. Prior to crew compartment entry, materials and equipment must be checked by an orbiter integrity clerk stationed outside the white-room entrance for compliance to program requirements. Analysis and source identification of crew compartment debris studies have been going on for two years. The objective of these studies is to determine and identify particulate generating materials and activities in the crew compartment. Results show a wide spectrum of many different types of materials. When source identification is made, corrective action is implemented to minimize or curtail further contaminate generation.

Improved solution accuracy for Landsat-4 (TDRSS-user) orbit determination

NASA Technical Reports Server (NTRS)

Oza, D. H.; Niklewski, D. J.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1994-01-01

This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using a Prototype Filter Smoother (PFS), with the accuracy of an established batch-least-squares system, the Goddard Trajectory Determination System (GTDS). The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and convariances for the sequential case) of solutions produced by the batch and sequential methods. The filtered and smoothed PFS orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 15 meters.

Astrodynamics 1991; Proceedings of the AAS/AIAA Astrodynamics Conference, Durango, CO, Aug. 19-22, 1991. Pts. 1, 2, and 3

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kaufman, B.; Alfriend, K.T.; Roehrich, R.L.

1992-01-01

The present conference on astrodynamics and advances in the astronautical sciences encompasses orbit determination, orbital debris, flexible-body dynamics and control, attitude dynamics and control, and topics related to the projects of the European space program. Specific issues addressed include a numerical approach to the angles-only initial orbit determination problem, precise orbit determination of the SPOT platform with DORIS, space-debris measurement and modeling, H(infinity)-optimized broadband compensator for wave-absorbing control, and the application of linear actuators for for telescope pointing control. Also addressed are attitude determination and dynamical performance in free drift for the Space Station Freedom, a Kalman filter for amore » gravity-gradient satellite, the positioning of the Eutelsat II satellite from supersynchronous transfer orbit to reduce satellite velocity-correction requirements, and trajectory analysis and issues.« less

Autonomous orbital navigation using Kepler's equation

NASA Technical Reports Server (NTRS)

Boltz, F. W.

1974-01-01

A simple method of determining the six elements of elliptic satellite orbits has been developed for use aboard manned and unmanned spacecraft orbiting the earth, moon, or any planet. The system requires the use of a horizon sensor or other device for determining the local vertical, a precision clock or timing device, and Apollo-type navigation equipment including an inertial measurement unit (IMU), a digital computer, and a coupling data unit. The three elements defining the in-plane motion are obtained from simultaneous measurements of central angle traversed around the planet and elapsed flight time using a linearization of Kepler's equation about a reference orbit. It is shown how Kalman filter theory may also be used to determine the in-plane orbital elements. The three elements defining the orbit orientation are obtained from position angles in celestial coordinates derived from the IMU with the spacecraft vertically oriented after alignment of the IMU to a known inertial coordinate frame.

Orbit Determination of the SELENE Satellites Using Multi-Satellite Data Types and Evaluation of SELENE Gravity Field Models

NASA Technical Reports Server (NTRS)

Goossens, S.; Matsumoto, K.; Noda, H.; Araki, H.; Rowlands, D. D.; Lemoine, F. G.

2011-01-01

The SELENE mission, consisting of three separate satellites that use different terrestrial-based tracking systems, presents a unique opportunity to evaluate the contribution of these tracking systems to orbit determination precision. The tracking data consist of four-way Doppler between the main orbiter and one of the two sub-satellites while the former is over the far side, and of same-beam differential VLBI tracking between the two sub-satellites. Laser altimeter data are also used for orbit determination. The contribution to orbit precision of these different data types is investigated through orbit overlap analysis. It is shown that using four-way and VLBI data improves orbit consistency for all satellites involved by reducing peak values in orbit overlap differences that exist when only standard two-way Doppler and range data are used. Including laser altimeter data improves the orbit precision of the SELENE main satellite further, resulting in very smooth total orbit errors at an average level of 18m. The multi-satellite data have also resulted in improved lunar gravity field models, which are assessed through orbit overlap analysis using Lunar Prospector tracking data. Improvements over a pre-SELENE model are shown to be mostly in the along-track and cross-track directions. Orbit overlap differences are at a level between 13 and 21 m with the SELENE models, depending on whether l-day data overlaps or I-day predictions are used.

Lunar Gravity Field Determination Using SELENE Same-Beam Differential VLBI Tracking Data

NASA Technical Reports Server (NTRS)

Goossens, S.; Matsumoto, K.; Liu, Q.; Kikuchi, F.; Sato, K.; Hanada, H.; Ishihara, Y.; Noda, H.; Kawano, N.; Namiki, N.;

Orbital evolution of space debris due to aerodynamic forces

NASA Astrophysics Data System (ADS)

Crowther, R.

1993-08-01

The concepts used in the AUDIT (Assessment Using Debris Impact Theory) debris modelling suite are introduced. A sensitivity analysis is carried out to determine the dominant parameters in the modelling process. A test case simulating the explosion of a satellite suggest that at the parent altitude there is a greater probability of collision with more massive fragments.

Evaluation of semiempirical atmospheric density models for orbit determination applications

NASA Technical Reports Server (NTRS)

Cox, C. M.; Feiertag, R. J.; Oza, D. H.; Doll, C. E.

1994-01-01

This paper presents the results of an investigation of the orbit determination performance of the Jacchia-Roberts (JR), mass spectrometer incoherent scatter 1986 (MSIS-86), and drag temperature model (DTM) atmospheric density models. Evaluation of the models was performed to assess the modeling of the total atmospheric density. This study was made generic by using six spacecraft and selecting time periods of study representative of all portions of the 11-year cycle. Performance of the models was measured for multiple spacecraft, representing a selection of orbit geometries from near-equatorial to polar inclinations and altitudes from 400 kilometers to 900 kilometers. The orbit geometries represent typical low earth-orbiting spacecraft supported by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). The best available modeling and orbit determination techniques using the Goddard Trajectory Determination System (GTDS) were employed to minimize the effects of modeling errors. The latest geopotential model available during the analysis, the Goddard earth model-T3 (GEM-T3), was employed to minimize geopotential model error effects on the drag estimation. Improved-accuracy techniques identified for TOPEX/Poseidon orbit determination analysis were used to improve the Tracking and Data Relay Satellite System (TDRSS)-based orbit determination used for most of the spacecraft chosen for this analysis. This paper shows that during periods of relatively quiet solar flux and geomagnetic activity near the solar minimum, the choice of atmospheric density model used for orbit determination is relatively inconsequential. During typical solar flux conditions near the solar maximum, the differences between the JR, DTM, and MSIS-86 models begin to become apparent. Time periods of extreme solar activity, those in which the daily and 81-day mean solar flux are high and change rapidly, result in significant differences between the models. During periods of high geomagnetic activity, the standard JR model was outperformed by DTM. Modification of the JR model to use a geomagnetic heating delay of 3 hours, as used in DTM, instead of the 6.7-hour delay produced results comparable to or better than the DTM performance, reducing definitive orbit solution ephermeris overlap differences by 30 to 50 percent. The reduction in the overlap differences would be useful for mitigating the impact of geomagnetic storms on orbit prediction.

Computer image processing - The Viking experience. [digital enhancement techniques

NASA Technical Reports Server (NTRS)

Green, W. B.

1977-01-01

Computer processing of digital imagery from the Viking mission to Mars is discussed, with attention given to subjective enhancement and quantitative processing. Contrast stretching and high-pass filtering techniques of subjective enhancement are described; algorithms developed to determine optimal stretch and filtering parameters are also mentioned. In addition, geometric transformations to rectify the distortion of shapes in the field of view and to alter the apparent viewpoint of the image are considered. Perhaps the most difficult problem in quantitative processing of Viking imagery was the production of accurate color representations of Orbiter and Lander camera images.

A Geometric Analysis to Protect Manned Assets from Newly Launched Objects - COLA Gap Analysis

NASA Technical Reports Server (NTRS)

Hametz, Mark E.; Beaver, Brian A.

2012-01-01

A safety risk was identified for the International Space Station (ISS) by The Aerospace Corporation following the launch of GPS IIR-20 (March 24, 2009), when the spent upper stage of the launch vehicle unexpectedly crossed inside the ISS notification box shortly after launch. This event highlighted a 56-hour vulnerability period following the end of the launch Collision Avoidance (COLA) process where the ISS would be unable to react to a conjunction with a newly launched object. Current launch COLA processes screen each launched object across the launch window to determine if an object's nominal trajectory is predicted to pass within 200 km of the ISS (or any other manned/mannable object), resulting in a launch time closure. These launch COLA screens are performed from launch through separation plus I 00 minutes. Once the objects are in orbit, they are cataloged and evaluated as part of routine on-orbit conjunction assessment processes. However, as the GPS IIR-20 scenario illustrated, there is a vulnerability period in the time line between the end of launch COLA coverage and the beginning of standard on-orbit COLA assessment activities. The gap between existing launch and on-orbit COLA processes is driven by the time it takes to track and catalog a launched object, identify a conjunction, and plan and execute a collision avoidance maneuver. For the ISS, the total time required to accomplish an of these steps is 56 hours. To protect human lives, NASA/JSC has requested that an US launches take additional steps to protect the ISS during this "COLA gap" period. The uncertainty in the state of a spent upper stage can be quite large after all bums are complete and all remaining propellants are expelled to safe the stage. Simply extending the launch COLA process an additional 56 hours is not a viable option as the 3-sigma position uncertainty will far exceed the 200 km miss-distance criterion. Additionally, performing a probability of collision (Pc) analysis over this period is also not practical due to the limiting effects of these large orbit state uncertainties. An estimated upper bound for Pc for a typical spent upper stage if nominally aligned for a direct broadside collision with the ISS is only on the order of 10-6. For a smaller manned object such as a Soyuz capsule, the risk level decreases to an order of 10'8 . In comparison, the Air Force Range policy (AFI 91-217) for launch COLAs would only eliminate launch opportunities when conjunctions with objects exceed a Pc of 10'5 This paper demonstrates a conservative geometry-based methodology that may be used to determine if launch opportunities pose a threat to the ISS during the COLA gap period. The NASA Launch Services Program at Kennedy Space Center has developed this COLA gap analysis method and employed it fQr three NASA missions to identify potential ISS conjunctions and corresponding launch window closures during the 56-hour at-risk period. In the analysis, for each launch opportunity, the nominal trajectory of the spent upper stage and the orbit state of the ISS are propagated over the 56 hour period. Each time the upper stage crosses the orbit plane of the ISS, the relative radial and argument of latitude separations are calculated. A window cutout is identified if these separation differences fall within a mission-specific violation box, which is determined from the evaluation of a Monte Carlo dispersions analysis that quantifies the potential variation in the upper stage radial and argument of latitude differences. This paper details the results of these analyses and their impacts to each mission.

Overall view of the Orbiter Servicing Structure within the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Overall view of the Orbiter Servicing Structure within the Orbiter Processing Facility at Kennedy Space Center. Can you see any hint of the Orbiter Discovery? It is in there. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Thermospheric density and wind retrieval from Swarm observations

NASA Astrophysics Data System (ADS)

Visser, Pieter; Doornbos, Eelco; van den IJssel, Jose; Teixeira da Encarnação, João

2013-11-01

The three-satellite ESA Swarm mission aims at mapping the Earth's global geomagnetic field at unprecedented spatial and temporal resolution and precision. Swarm also aims at observing thermospheric density and possibly horizontal winds. Precise orbit determination (POD) and Thermospheric Density and Wind (TDW) chains form part of the Swarm Constellation and Application Facility (SCARF), which will provide the so-called Level 2 products. The POD and TDW chains generate the orbit, accelerometer calibration, and thermospheric density and wind Level 2 products. The POD and TDW chains have been tested with data from the CHAMP and GRACE missions, indicating that a 3D orbit precision of about 10 cm can be reached. In addition, POD allows to determine daily accelerometer bias and scale factor values with a precision of around 10-15 nm/s2 and 0.01-0.02, respectively, for the flight direction. With these accelerometer calibration parameter values, derived thermospheric density is consistent at the 9-11% level (standard deviation) with values predicted by models (taking into account that model values are 20-30% higher). The retrieval of crosswinds forms part of the processing chain, but will be challenging. The Swarm observations will be used for further developing and improving density and wind retrieval algorithms.

Independent Orbiter Assessment (IOA): Analysis of the hydraulics/water spray boiler subsystem

NASA Technical Reports Server (NTRS)

Duval, J. D.; Davidson, W. R.; Parkman, William E.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items (PCIs). To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Orbiter Hydraulics/Water Spray Boiler Subsystem. The hydraulic system provides hydraulic power to gimbal the main engines, actuate the main engine propellant control valves, move the aerodynamic flight control surfaces, lower the landing gear, apply wheel brakes, steer the nosewheel, and dampen the external tank (ET) separation. Each hydraulic system has an associated water spray boiler which is used to cool the hydraulic fluid and APU lubricating oil. The IOA analysis process utilized available HYD/WSB hardware drawings, schematics and documents for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 430 failure modes analyzed, 166 were determined to be PCIs.

Independent Orbiter Assessment (IOA): Analysis of the remote manipulator system

NASA Technical Reports Server (NTRS)

Tangorra, F.; Grasmeder, R. F.; Montgomery, A. D.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items (PCIs). To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Orbiter Remote Manipulator System (RMS) are documented. The RMS hardware and software are primarily required for deploying and/or retrieving up to five payloads during a single mission, capture and retrieve free-flying payloads, and for performing Manipulator Foot Restraint operations. Specifically, the RMS hardware consists of the following components: end effector; displays and controls; manipulator controller interface unit; arm based electronics; and the arm. The IOA analysis process utilized available RMS hardware drawings, schematics and documents for defining hardware assemblies, components and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 574 failure modes analyzed, 413 were determined to be PCIs.

STV fueling options

NASA Technical Reports Server (NTRS)

Flemming, Ken

1991-01-01

Lunar vehicles that will be space based and reusable will require resupply of propellants in orbit. Approximately 75 pct. of the total mass delivered to low earth orbit will be propellants. Consequently, the propellant management techniques selected for Space Exploration Initiative (SEI) orbital operations will have a major influence on the overall SEI architecture. Five proposed propellant management facility (PMF) concepts were analyzed and compared in order to determine the best method of resupplying reusable, space based Lunar Transfer Vehicles (LTVs). The processing time needed at the Space Station to prepare LTV for its next lunar mission was estimated for each of the PMF concepts. The estimated times required to assemble and maintain the different PMF concepts were also compared. The results of the maintenance analysis were similar, with co-orbiting depots needing 100 to 350 pct. more annual maintenance. The first few external tanks mating operations at KSC encountered many problems that could cause serious lunar mission schedule delays. The use of drop tanks on lunar vehicles increases by a factor of four the number of critical propellant interface disturbances.

Independent Orbiter Assessment (IOA): Analysis of the electrical power generation/fuel cell powerplant subsystem

NASA Technical Reports Server (NTRS)

Brown, K. L.; Bertsch, P. J.

1986-01-01

Results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Electrical Power Generation (EPG)/Fuel Cell Powerplant (FCP) hardware. The EPG/FCP hardware is required for performing functions of electrical power generation and product water distribution in the Orbiter. Specifically, the EPG/FCP hardware consists of the following divisions: (1) Power Section Assembly (PSA); (2) Reactant Control Subsystem (RCS); (3) Thermal Control Subsystem (TCS); and (4) Water Removal Subsystem (WRS). The IOA analysis process utilized available EPG/FCP hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Geographical representation of radial orbit perturbations due to ocean tides: Implications for satellite altimetry

NASA Technical Reports Server (NTRS)

Bettadpur, Srinivas V.; Eanes, Richard J.

1994-01-01

In analogy to the geographical representation of the zeroth-order radial orbit perturbations due to the static geopotential, similar relationships have been derived for radial orbit perturbations due to the ocean tides. At each location these perturbations are seen to be coherent with the tide height variations. The study of this singularity is of obvious importance to the estimation of ocean tides from satellite altimeter data. We derive analytical expressions for the sensitivity of altimeter derived ocean tide models to the ocean tide force model induced errors in the orbits of the altimeter satellite. In particular, we focus on characterizing and quantifying the nonresonant tidal orbit perturbations, which cannot be adjusted into the empirical accelerations or radial perturbation adjustments commonly used during orbit determination and in altimeter data processing. As an illustration of the utility of this technique, we study the differences between a TOPEX/POSEIDON-derived ocean tide model and the Cartwright and Ray 1991 Geosat model. This analysis shows that nearly 60% of the variance of this difference for M(sub 2) can be explained by the Geosat radial orbit eror due to the omission of coefficients from the GEM-T2 background ocean tide model. For O(sub 1), K(sub 1), S(sub 2), and K(sub 2) the orbital effects account for approximately 10 to 40% of the variances of these differences. The utility of this technique to assessment of the ocean tide induced errors in the TOPEX/POSEIDON-derived tide models is also discussed.

Study of a homotopy continuation method for early orbit determination with the Tracking and Data Relay Satellite System (TDRSS)

NASA Technical Reports Server (NTRS)

Smith, R. L.; Huang, C.

1986-01-01

A recent mathematical technique for solving systems of equations is applied in a very general way to the orbit determination problem. The study of this technique, the homotopy continuation method, was motivated by the possible need to perform early orbit determination with the Tracking and Data Relay Satellite System (TDRSS), using range and Doppler tracking alone. Basically, a set of six tracking observations is continuously transformed from a set with known solution to the given set of observations with unknown solutions, and the corresponding orbit state vector is followed from the a priori estimate to the solutions. A numerical algorithm for following the state vector is developed and described in detail. Numerical examples using both real and simulated TDRSS tracking are given. A prototype early orbit determination algorithm for possible use in TDRSS orbit operations was extensively tested, and the results are described. Preliminary studies of two extensions of the method are discussed: generalization to a least-squares formulation and generalization to an exhaustive global method.

Applicability of meteor radiant determination methods depending on orbit type. II. Low-eccentric orbits

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.

General view from inside the payload bay or the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view from inside the payload bay or the Orbiter Discovery approximately along its centerline looking forward toward the bulkhead of the forward fuselage. Note the panels and insulation removed for access to the orbiter's subsystems for inspection and post-mission processing. Also note the airlock and the beam-truss attach structure supporting it and attaching it to the payload bay sill longerons. In this view the docking ring and airlock hatches have been removed. This photo was taken during the processing of the Orbiter Discovery after its final mission and in preparation for its transition to the National Air and Space Museum. This view was taken in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Spaceflight mechanics 1992; Proceedings of the 2nd AAS/AIAA Meeting, Colorado Springs, CO, Feb. 24-26, 1992. Pts. 1 & 2

NASA Astrophysics Data System (ADS)

Diehl, Roger E.; Schinnerer, Ralph G.; Williamson, Walton E.; Boden, Daryl G.

The present conference discusses topics in orbit determination, tethered satellite systems, celestial mechanics, guidance optimization, flexible body dynamics and control, attitude dynamics and control, Mars mission analyses, earth-orbiting mission analysis/debris, space probe mission analyses, and orbital computation numerical analyses. Attention is given to electrodynamic forces for control of tethered satellite systems, orbiting debris threats to asteroid flyby missions, launch velocity requirements for interceptors of short range ballistic missiles, transfers between libration-point orbits in the elliptic restricted problem, minimum fuel spacecraft reorientation, orbital guidance for hitting a fixed point at maximum speed, efficient computation of satellite visibility periods, orbit decay and reentry prediction for space debris, and the determination of satellite close approaches.

Spaceflight mechanics 1992; Proceedings of the 2nd AAS/AIAA Meeting, Colorado Springs, CO, Feb. 24-26, 1992. Pts. 1 & 2

NASA Technical Reports Server (NTRS)

Diehl, Roger E. (Editor); Schinnerer, Ralph G. (Editor); Williamson, Walton E. (Editor); Boden, Daryl G. (Editor)

1992-01-01

The present conference discusses topics in orbit determination, tethered satellite systems, celestial mechanics, guidance optimization, flexible body dynamics and control, attitude dynamics and control, Mars mission analyses, earth-orbiting mission analysis/debris, space probe mission analyses, and orbital computation numerical analyses. Attention is given to electrodynamic forces for control of tethered satellite systems, orbiting debris threats to asteroid flyby missions, launch velocity requirements for interceptors of short range ballistic missiles, transfers between libration-point orbits in the elliptic restricted problem, minimum fuel spacecraft reorientation, orbital guidance for hitting a fixed point at maximum speed, efficient computation of satellite visibility periods, orbit decay and reentry prediction for space debris, and the determination of satellite close approaches.

FORMOSAT-3/COSMIC POD Data Processing and Initial Results

NASA Astrophysics Data System (ADS)

Tang, C.

2006-12-01

The six satellites of the collaborative Taiwan-U.S. FORMOSAT-3/COSMIC (Constellation Observing System for Meteorology, Ionosphere, and Climate) space program were successfully launched from Vandenberg, U.S.A. on April 15, 2006. As of September 7, 2006, one satellite (FM5) has already been transferred to the 800-km final orbit, while the other five satellites (FM1-4 and FM6) are currently waiting in the ~520-km parking orbit for subsequent orbit raising deployment. There are two GPS antennas with different orientation onboard each satellite whose measurements are used specifically for precise orbit determination (POD). The received GPS signals by the POD antennas were rather sparse and unstable in the initial 5 weeks. Since then, the available GPS measurements have gradually increased from 10-20% in the early stage to almost 90% in 11 weeks after the launch. For the two POD antennas (POD+X and POD-X), one antenna can perform normally and record observations from up to 9 GPS satellites in view; however, the other antenna is programmed to track up to 4 GPS satellites due to onboard memory limitation. For this reason, we first performed orbit computation using zero-difference GPS phases collected by the normal antenna. For each day's orbit computation, we designed a 6-hr (25%) overlap for inner orbital accuracy assessment, and overlap analysis shows that the achievable 3D RMS was around 19 cm, or 11 cm per axis. In a separate effort, orbit computation based on the lesser antenna was also performed. The orbital difference between the results obtained from the two antennas was significant, with a 3D RMS value of 64 cm. The early results indicate that more work is needed in order to incorporate GPS data from both antennas into a unified solution.

KOI2138 -- a Spin-Orbit Aligned Intermediate Period Super-Earth

NASA Astrophysics Data System (ADS)

Barnes, Jason W.

2015-11-01

A planet's formation and evolution are encoded in spin-orbit alignment -- the planet's inclination relative to its star's equatorial plane. While the solar system's spin-orbit aligned planets indicate our own relatively quiescent history, many close-in giant planets show significant misalignment. Some planets even orbit retrograde! Hot Jupiters, then, have experienced fundamentally different histories than we experienced here in the solar system. In this presentation, I will show a new determination of the spin-orbit alignment of 2.1 REarth exoplanet candidate KOI2138. KOI2138 shows a gravity-darkened transit lightcurve that is consistent with spin-orbit alignment. This measurement is important because the only other super-Earth with an alignment determination (55 Cnc e, orbit period 0.74 days) is misaligned. With an orbital period of 23.55 days, KOI2138 is far enough from its star to avoid tidal orbit evolution. Therefore its orbit is likely primordial, and hence it may represent the tip of an iceberg of terrestrial, spin-orbit aligned planets that have histories that more closely resemble that of the solar system's terrestrial planets.

Mission Design for the Lunar Reconnaissance Orbiter

NASA Technical Reports Server (NTRS)

Beckman, Mark

2007-01-01

The Lunar Reconnaissance Orbiter (LRO) will be the first mission under NASA's Vision for Space Exploration. LRO will fly in a low 50 km mean altitude lunar polar orbit. LRO will utilize a direct minimum energy lunar transfer and have a launch window of three days every two weeks. The launch window is defined by lunar orbit beta angle at times of extreme lighting conditions. This paper will define the LRO launch window and the science and engineering constraints that drive it. After lunar orbit insertion, LRO will be placed into a commissioning orbit for up to 60 days. This commissioning orbit will be a low altitude quasi-frozen orbit that minimizes stationkeeping costs during commissioning phase. LRO will use a repeating stationkeeping cycle with a pair of maneuvers every lunar sidereal period. The stationkeeping algorithm will bound LRO altitude, maintain ground station contact during maneuvers, and equally distribute periselene between northern and southern hemispheres. Orbit determination for LRO will be at the 50 m level with updated lunar gravity models. This paper will address the quasi-frozen orbit design, stationkeeping algorithms and low lunar orbit determination.

Mechanisms controlling the complete accretionary beach state sequence

NASA Astrophysics Data System (ADS)

Dubarbier, Benjamin; Castelle, Bruno; Ruessink, Gerben; Marieu, Vincent

2017-06-01

Accretionary downstate beach sequence is a key element of observed nearshore morphological variability along sandy coasts. We present and analyze the first numerical simulation of such a sequence using a process-based morphodynamic model that solves the coupling between waves, depth-integrated currents, and sediment transport. The simulation evolves from an alongshore uniform barred beach (storm profile) to an almost featureless shore-welded terrace (summer profile) through the highly alongshore variable detached crescentic bar and transverse bar/rip system states. A global analysis of the full sequence allows determining the varying contributions of the different hydro-sedimentary processes. Sediment transport driven by orbital velocity skewness is critical to the overall onshore sandbar migration, while gravitational downslope sediment transport acts as a damping term inhibiting further channel growth enforced by rip flow circulation. Accurate morphological diffusivity and inclusion of orbital velocity skewness opens new perspectives in terms of morphodynamic modeling of real beaches.

Implementation of Altimetry Data in the GIPSY POD Software Package

NASA Technical Reports Server (NTRS)

Stauch, Jason R.; Gold, Kenn; Born, George H.

2001-01-01

Altimetry data has been used extensively to acquire data about characteristics of the Earth, the Moon, and Mars. More recently, the idea of using altimetry for orbit determination has also been explored. This report discusses modifications to JPL's GIPSY/OASIS II software to include altimetry data as an observation type for precise orbit determination. The mathematical foundation of using altimetry for the purpose of orbit determination is presented, along with results.

An Integrated Vision-Based System for Spacecraft Attitude and Topology Determination for Formation Flight Missions

NASA Technical Reports Server (NTRS)

Rogers, Aaron; Anderson, Kalle; Mracek, Anna; Zenick, Ray

2004-01-01

With the space industry's increasing focus upon multi-spacecraft formation flight missions, the ability to precisely determine system topology and the orientation of member spacecraft relative to both inertial space and each other is becoming a critical design requirement. Topology determination in satellite systems has traditionally made use of GPS or ground uplink position data for low Earth orbits, or, alternatively, inter-satellite ranging between all formation pairs. While these techniques work, they are not ideal for extension to interplanetary missions or to large fleets of decentralized, mixed-function spacecraft. The Vision-Based Attitude and Formation Determination System (VBAFDS) represents a novel solution to both the navigation and topology determination problems with an integrated approach that combines a miniature star tracker with a suite of robust processing algorithms. By combining a single range measurement with vision data to resolve complete system topology, the VBAFDS design represents a simple, resource-efficient solution that is not constrained to certain Earth orbits or formation geometries. In this paper, analysis and design of the VBAFDS integrated guidance, navigation and control (GN&C) technology will be discussed, including hardware requirements, algorithm development, and simulation results in the context of potential mission applications.

General view of the aft fuselage of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the aft fuselage of the Orbiter Discovery looking forward showing Space Shuttle Main Engines (SSMEs) installed in positions one and three and an SSME on the process of being installed in position two. This photograph was taken in the Orbiter Processing Facility at the Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Initial Results from On-Orbit Testing of the Fram Memory Test Experiment on the Fastsat Micro-Satellite

NASA Technical Reports Server (NTRS)

MacLeond, Todd C.; Sims, W. Herb; Varnavas,Kosta A.; Ho, Fat D.

2011-01-01

The Memory Test Experiment is a space test of a ferroelectric memory device on a low Earth orbit satellite that launched in November 2010. The memory device being tested is a commercial Ramtron Inc. 512K memory device. The circuit was designed into the satellite avionics and is not used to control the satellite. The test consists of writing and reading data with the ferroelectric based memory device. Any errors are detected and are stored on board the satellite. The data is sent to the ground through telemetry once a day. Analysis of the data can determine the kind of error that was found and will lead to a better understanding of the effects of space radiation on memory systems. The test is one of the first flight demonstrations of ferroelectric memory in a near polar orbit which allows testing in a varied radiation environment. The initial data from the test is presented. This paper details the goals and purpose of this experiment as well as the development process. The process for analyzing the data to gain the maximum understanding of the performance of the ferroelectric memory device is detailed.

Modification of an impulse-factoring orbital transfer technique to account for orbit determination and maneuver execution errors

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.

General view of the Orbiter Discovery in the Orbiter Processing ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the Orbiter Discovery in the Orbiter Processing Facility at Kennedy Space Center showing the payload bay doors open exposing the heat-dissipating radiator panels located on the inside of the payload bay doors. Also in the view is the boom portion of the boom sensor system deployed as part of the return to flight procedures after STS-107 to inspect the orbiter's thermal protection system. The Remote Manipulator System, the "Canadarm", and the airlock are seen in the background of the image. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility oversee installation of the liquid oxygen feedline for the 17-inch disconnect on the orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

NASA Image and Video Library

2003-11-11

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility oversee installation of the liquid oxygen feedline for the 17-inch disconnect on the orbiter Discovery. The 17-inch liquid oxygen and liquid hydrogen disconnects provide the propellant feed interface from the external tank to the orbiter main propulsion system and the three Shuttle main engines.

The limits of direct satellite tracking with the Global Positioning System (GPS)

NASA Technical Reports Server (NTRS)

Bertiger, W. I.; Yunck, T. P.

1988-01-01

Recent advances in high precision differential Global Positioning System-based satellite tracking can be applied to the more conventional direct tracking of low earth satellites. To properly evaluate the limiting accuracy of direct GPS-based tracking, it is necessary to account for the correlations between the a-priori errors in GPS states, Y-bias, and solar pressure parameters. These can be obtained by careful analysis of the GPS orbit determination process. The analysis indicates that sub-meter accuracy can be readily achieved for a user above 1000 km altitude, even when the user solution is obtained with data taken 12 hours after the data used in the GPS orbit solutions.

Infrared Spectroscopy of Symbiotic Stars. XI. Orbits for Southern S-type Systems: Hen 3-461, SY Mus, Hen 3-828, AND AR Pav

NASA Astrophysics Data System (ADS)

Fekel, Francis C.; Hinkle, Kenneth H.; Joyce, Richard R.; Wood, Peter R.

2017-01-01

Employing new infrared radial velocities, we have computed spectroscopic orbits of the cool giants in four southern S-type symbiotic systems. The orbits for two of the systems, Hen 3-461 and Hen 3-828, have been determined for the first time, while orbits of the other two, SY Mus and AR Pav, have previously been determined. For the latter two systems, we compare our results with those in the literature. The low mass of the secondary of SY Mus suggests that it has gone through a common envelope phase. Hen 3-461 has an orbital period of 2271 days, one of the longest currently known for S-type symbiotic systems. That period is very different from the orbital period proposed previously from its photometric variations. The other three binaries have periods between 600 and 700 day, values that are typical for S-type symbiotic orbits. Basic properties of the M giant components and the distance to each system are determined.

Angles-only relative orbit determination in low earth orbit

NASA Astrophysics Data System (ADS)

Ardaens, Jean-Sébastien; Gaias, Gabriella

2018-06-01

The paper provides an overview of the angles-only relative orbit determination activities conducted to support the Autonomous Vision Approach Navigation and Target Identification (AVANTI) experiment. This in-orbit endeavor was carried out by the German Space Operations Center (DLR/GSOC) in autumn 2016 to demonstrate the capability to perform spaceborne autonomous close-proximity operations using solely line-of-sight measurements. The images collected onboard have been reprocessed by an independent on-ground facility for precise relative orbit determination, which served as ultimate instance to monitor the formation safety and to characterize the onboard navigation and control performances. During two months, several rendezvous have been executed, generating a valuable collection of images taken at distances ranging from 50 km to only 50 m. Despite challenging experimental conditions characterized by a poor visibility and strong orbit perturbations, angles-only relative positioning products could be continuously derived throughout the whole experiment timeline, promising accuracy at the meter level during the close approaches. The results presented in the paper are complemented with former angles-only experience gained with the PRISMA satellites to better highlight the specificities induced by different orbits and satellite designs.

INFRARED SPECTROSCOPY OF SYMBIOTIC STARS. XI. ORBITS FOR SOUTHERN S-TYPE SYSTEMS: HEN 3-461, SY MUS, HEN 3-828, AND AR PAV

DOE Office of Scientific and Technical Information (OSTI.GOV)

Fekel, Francis C.; Hinkle, Kenneth H.; Joyce, Richard R.

Employing new infrared radial velocities, we have computed spectroscopic orbits of the cool giants in four southern S-type symbiotic systems. The orbits for two of the systems, Hen 3-461 and Hen 3-828, have been determined for the first time, while orbits of the other two, SY Mus and AR Pav, have previously been determined. For the latter two systems, we compare our results with those in the literature. The low mass of the secondary of SY Mus suggests that it has gone through a common envelope phase. Hen 3-461 has an orbital period of 2271 days, one of the longest currently known for S-type symbiotic systems.more » That period is very different from the orbital period proposed previously from its photometric variations. The other three binaries have periods between 600 and 700 day, values that are typical for S-type symbiotic orbits. Basic properties of the M giant components and the distance to each system are determined.« less

Contamination control and cleanliness level integrity for the Space Shuttle Orbiter PLB, payloads and facilities at KSC

NASA Technical Reports Server (NTRS)

Bartelson, D.

1984-01-01

The PLB, its cargo, and payload canister must satisfy the cleanliness requirements of visual clean (VC) level 1, 2, 3, or special as stated in NASA document SN-C-0005A. The specific level of cleanliness is chosen by the payload bay customer for their mission. During orbiter turnaround processing at KSC, the payload bay is exposed to the environments of the Orbiter Processing Facility (OPF) and the Payload Changeout Room (PCR). In supportive response to the orbiter payload bay/facility interface, it is necessary that the facility environment be controlled and monitored to protect the cleanliness/environmental integrity of the payload bay and its cargo. Techniques used to meet environmental requirements during orbiter processing are introduced.

Phase change paint tests on Rockwell orbiter/tank and orbiter alone configurations (OH3A/OH3B)

NASA Technical Reports Server (NTRS)

Quan, M.; Craig, C.

1974-01-01

Wind tunnel tests were conducted on scale models of the space shuttle orbiter and external tank. The tests were designed to determine the basic heating rate and interference effects on the orbiter-tank configuration and to analyze the effectiveness of the thermal protective system on the reentry vehicle. The phase change paint techniques were used to determine areodynamic heating rates. Oil flow and schlieren photographs were used for flow visualization.

Determining characteristics of artificial near-Earth objects using observability analysis

NASA Astrophysics Data System (ADS)

Friedman, Alex M.; Frueh, Carolin

2018-03-01

Observability analysis is a method for determining whether a chosen state of a system can be determined from the output or measurements. Knowledge of state information availability resulting from observability analysis leads to improved sensor tasking for observation of orbital debris and better control of active spacecraft. This research performs numerical observability analysis of artificial near-Earth objects. Analysis of linearization methods and state transition matrices is performed to determine the viability of applying linear observability methods to the nonlinear orbit problem. Furthermore, pre-whitening is implemented to reformulate classical observability analysis. In addition, the state in observability analysis is typically composed of position and velocity; however, including object characteristics beyond position and velocity can be crucial for precise orbit propagation. For example, solar radiation pressure has a significant impact on the orbit of high area-to-mass ratio objects in geosynchronous orbit. Therefore, determining the time required for solar radiation pressure parameters to become observable is important for understanding debris objects. In order to compare observability analysis results with and without measurement noise and an extended state, quantitative measures of observability are investigated and implemented.

Flight Mechanics/Estimation Theory Symposium

NASA Technical Reports Server (NTRS)

1978-01-01

Satellite attitude determination and control, orbit determination, and onboard and ground attitude determination procedures are among the topics discussed. Other topics covered include: effect of atmosphere on Venus orbiter navigation; satellite-to-satellite tracking; and satellite onboard navigation using global positioning system data.

The long-term dynamical behavior of short-period comets

NASA Technical Reports Server (NTRS)

Levison, Harold F.; Duncan, Martin J.

1993-01-01

The orbits of the known short-period comets under the influence of the Sun and all the planets except Mercury and Pluto are numerically integrated. The calculation was undertaken in order to determine the dynamical lifetimes for these objects as well as explaining the current orbital element distribution. It is found that a comet can move between Jupiter-family and Halley-family comets several times in its dynamical lifetime. The median lifetime of the known short-period comets from the time they are first injected into a short-period comet orbit to ultimate ejection is approximately 50,000 years. The very flat inclination distribution of Jupiter-family comets is observed to become more distended as it ages. The only possible explanation for the observed flat distribution is that the comets become extinct before their inclination distribution can change significantly. It is shown that the anomalous concentration of the argument of perihelion of Jupiter-family comets near 0 and 180 deg is a direct result of their aphelion distance being close to 5.2AU and the comet being recently perturbed onto a Jupiter-family orbit. Also the concentration of their aphelion near Jupiter's orbit is a result of the conservation of the Tisserand invariant during the capture process.

Precise Orbit Determination Of Low Earth Satellites At AIUB Using GPS And SLR Data

NASA Astrophysics Data System (ADS)

Jaggi, A.; Bock, H.; Thaller, D.; Sosnica, K.; Meyer, U.; Baumann, C.; Dach, R.

2013-12-01

An ever increasing number of low Earth orbiting (LEO) satellites is, or will be, equipped with retro-reflectors for Satellite Laser Ranging (SLR) and on-board receivers to collect observations from Global Navigation Satellite Systems (GNSS) such as the Global Positioning System (GPS) and the Russian GLONASS and the European Galileo systems in the future. At the Astronomical Institute of the University of Bern (AIUB) LEO precise orbit determination (POD) using either GPS or SLR data is performed for a wide range of applications for satellites at different altitudes. For this purpose the classical numerical integration techniques, as also used for dynamic orbit determination of satellites at high altitudes, are extended by pseudo-stochastic orbit modeling techniques to efficiently cope with potential force model deficiencies for satellites at low altitudes. Accuracies of better than 2 cm may be achieved by pseudo-stochastic orbit modeling for satellites at very low altitudes such as for the GPS-based POD of the Gravity field and steady-state Ocean Circulation Explorer (GOCE).

Decrease in the orbital period of Hercules X-1

NASA Technical Reports Server (NTRS)

Deeter, John E.; Boynton, Paul E.; Miyamoto, Sigenori; Kitamoto, Shunji; Nagase, Fumiaki; Kawai, Nobuyuki

1991-01-01

From a pulse-timing analysis of Ginga observations of the binary X-ray pulsar Her X-1 obtained during the interval 1989 April-June local orbital parameters are determined for a short high state. An orbital epoch is also determined in the adjacent main high state. By comparing these orbital solutions with previously published results, a decrease is detected in the orbital period for Her X-1 over the interval 1971-1989. The value is substantially larger than the value predicted from current estimates of the mass-transfer rate, and motivates consideration of other mechanisms of mass transfer and/or mass loss. A second result from these observations is a close agreement between orbital parameters determined separately in main high and short high states. This agreement places strong constraints on the obliquity of the stellar companion, HZ Her, if undergoing forced precession with a 35-day period. As a consequence further doubt is placed on the slaved-disk model as the underlying cause of the 35-day cycle in Her X-1.

Dawn Orbit Determination Team: Trajectory and Gravity Prediction Performance During Vesta Science Phases

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.

Research and Technology Report. Goddard Space Flight Center

NASA Technical Reports Server (NTRS)

Soffen, Gerald (Editor); Truszkowski, Walter (Editor); Ottenstein, Howard (Editor); Frost, Kenneth (Editor); Maran, Stephen (Editor); Walter, Lou (Editor); Brown, Mitch (Editor)

1996-01-01

This issue of Goddard Space Flight Center's annual report highlights the importance of mission operations and data systems covering mission planning and operations; TDRSS, positioning systems, and orbit determination; ground system and networks, hardware and software; data processing and analysis; and World Wide Web use. The report also includes flight projects, space sciences, Earth system science, and engineering and materials.

Astrometry and Geostationary Satellites in Venezuela

NASA Astrophysics Data System (ADS)

Lacruz, E.; Abad, C.

2015-10-01

We present the current status and the first results of the astrometric project CIDA - ABAE for tracking geo-stationary satellites. This project aims to determine a preliminary orbit for the Venezuelan satellite VENESAT-1, using astrometric positions obtained from an optical telescope. The results presented here are based on observations from the Luepa space tracking ground station in Venezuela, which were processed using astrometric procedures.

Radar Orbit Analysis Tool Using Least Squares Estimator

DTIC Science & Technology

2007-09-01

g gravity................................................................................................ km/s2 bodyg −2 v gravity due to two ...motion, it is necessary to determine how the dynamics between the two groups differ. One solution is to develop a model that can detect non...with just J2 and two -body terms was also addressed. Methodology Solving the estimation problem required dividing the process into four stages

Summary of the orbit determination of NOZOMI spacecraft for all the mission period

NASA Astrophysics Data System (ADS)

Yoshikawa, Makoto; Kawaguchi, Jun'Ichiro; Yamakawa, Hiroshi; Kato, Takaji; Ichikawa, Tsutomu; Ohnishi, Takafumi; Ishibashi, Shiro

2005-07-01

Japanese first Mars explorer NOZOMI, which was launched in July 1998, suffered several problems during the operation period of more than five years. It could have reached near Mars at the end of 2003, but it was not put into the orbit around Mars. Although NOZOMI was not able to execute its main mission, it provided us a lot of good experiences from the point of the orbit determination of spacecraft. One of the most difficult works was the orbit determination for the period without the telemetry. In this period, for the most of the time the high gain antenna did not point to the earth because of a constraint of the attitude. Therefore, the quality of the tracking data was not good, and for some period it was impossible to get the tracking data at all. Under such critical condition, we managed to get the solution of the orbit, and in a near-miraculous way, we were able to control NOZOMI and execute two earth swingbys successfully. Other issues related to the orbit determination are the spin modulation, the solar radiation pressure, the small force related to the attitude change, and the solar conjunction. We tried to solve these issues by the conventional way using range and Doppler data. However, we also tried the new method, that is the orbit determination by using the Delta-VLBI method (VLBI: Very Long Baseline Interferometry). In addition to this, we tried optical observations of NOZOMI at the earth swingbys.

Time Biases in laser ranging measurements; impacts on geodetic products (Reference Frame and Orbitography)

NASA Astrophysics Data System (ADS)

Belli, A.; Exertier, P.; Lemoine, F. G.; Chinn, D. S.; Zelensky, N. P.

2017-12-01

The GGOS objectives are to maintain a geodetic network with an accuracy of 1 mm and a stability of 0.1 mm per year. For years, the laser ranging technique, which provide very accurate absolute distances to geodetic targets enable to determine the scale factor as well as coordinates of the geocenter. In order to achieve this goal, systematic errors appearing in the laser ranging measurements must be considered and solved. In addition to Range Bias (RB), which is the primary source of uncertainty of the technique, Time Bias (TB) has been recently detected by using the Time Transfer by Laser Link (T2L2) space instrument capability on-board the satellite Jason-2. Instead of determining TB through the precise orbit determination that is applied to commonly used geodetic targets like LAGEOS to estimate global geodetic products, we have developed, independently, a dedicated method to transfer time between remote satellite laser ranging stations. As a result, the evolving clock phase shift to UTC of around 30 stations has been determined under the form of time series of time bias per station from 2008 to 2016 with an accuracy of 3-4 ns. It demonstrated the difficulty, in terms of Time & Frequency used technologies, to locally maintain accuracy and long term stability at least in the range of 100 ns that is the current requirement for time measurements (UTC) for the laser ranging technique. Because some laser ranging stations oftently exceed this limit (from 100 ns to a few μs) we have been studying these effects first on the precision orbit determination itself, second on the station positioning. We discuss the impact of TB on LAGEOS and Jason-2 orbits, which appears to affect the along-track component essentially. We also investigate the role of TB in global geodetic parameters as the station coordinates. Finally, we propose to provide the community with time series of time bias of laser ranging stations, under the form of a data- handling-file in order to be included in each orbit determination process that is using laser ranging data since 2008.

The challenge of precise orbit determination for STSAT-2C using extremely sparse SLR data

NASA Astrophysics Data System (ADS)

Kim, Young-Rok; Park, Eunseo; Kucharski, Daniel; Lim, Hyung-Chul; Kim, Byoungsoo

2016-03-01

The Science and Technology Satellite (STSAT)-2C is the first Korean satellite equipped with a laser retro-reflector array for satellite laser ranging (SLR). SLR is the only on-board tracking source for precise orbit determination (POD) of STSAT-2C. However, POD for the STSAT-2C is a challenging issue, as the laser measurements of the satellite are extremely sparse, largely due to the inaccurate two-line element (TLE)-based orbit predictions used by the SLR tracking stations. In this study, POD for the STSAT-2C using extremely sparse SLR data is successfully implemented, and new laser-based orbit predictions are obtained. The NASA/GSFC GEODYN II software and seven-day arcs are used for the SLR data processing of two years of normal points from March 2013 to May 2015. To compensate for the extremely sparse laser tracking, the number of estimation parameters are minimized, and only the atmospheric drag coefficients are estimated with various intervals. The POD results show that the weighted root mean square (RMS) post-fit residuals are less than 10 m, and the 3D day boundaries vary from 30 m to 3 km. The average four-day orbit overlaps are less than 20/330/20 m for the radial/along-track/cross-track components. The quality of the new laser-based prediction is verified by SLR observations, and the SLR residuals show better results than those of previous TLE-based predictions. This study demonstrates that POD for the STSAT-2C can be successfully achieved against extreme sparseness of SLR data, and the results can deliver more accurate predictions.

Spin-orbit-torque-induced skyrmion dynamics for different types of spin-orbit coupling

NASA Astrophysics Data System (ADS)

Lee, Seung-Jae; Kim, Kyoung-Whan; Lee, Hyun-Woo; Lee, Kyung-Jin

2018-06-01

We investigate current-induced skyrmion dynamics in the presence of Dzyaloshinskii-Moriya interaction and spin-orbit spin-transfer torque corresponding to various types of spin-orbit coupling. We determine the symmetries of Dzyaloshinskii-Moriya interaction and spin-orbit spin-transfer torque based on linear spin-orbit coupling model. We find that like interfacial Dzyaloshinskii-Moriya interaction (Rashba spin-orbit coupling) and bulk Dzyaloshinskii-Moriya interaction (Weyl spin-orbit coupling), Dresselhaus spin-orbit coupling also has a possibility for stabilizing skyrmion and current-induced skyrmion dynamics.

Common Orbital Infections ~ State of the Art ~ Part I

PubMed Central

Hamed-Azzam, Shirin; AlHashash, Islam; Briscoe, Daniel; Rose, Geoffrey E; Verity, David H.

2018-01-01

Infections of the orbit and periorbita are relatively frequent, and can cause significant local and systemic morbidity. Loss of vision occurs in more than 10% of patients, and systemic sequelae can include meningitis, intracranial abscess, and death. Numerous organisms infect the orbit, but the most common are bacteria. There are many methods through which orbital infections occur, with infection from the neighboring ethmoid sinuses the most likely cause for all age groups. Prompt management is essential in suspected orbital cellulitis, and involves urgent intravenous antibiotics, rehydration, and treatment of any co-existent underlying systemic disease, e.g., diabetes, renal failure. This review summarizes the common infectious processes of the orbit in both pediatric and adult groups. We review pathophysiology, symptoms, signs, and treatment for infectious orbital processes. PMID:29719647

Space Shuttle orbit determination using empirical force modeling of attitude maneuvers for the German MOMS-02/D2 mission

NASA Technical Reports Server (NTRS)

Vonbraun, C.; Reigber, Christoph

1994-01-01

In the spring of 1993, the MOMS-02 (modular Optoelectronic Multispectral Scanner) camera, as part of the second German Spacelab mission aboard STS-55, successfully took digital threefold stereo images of the surface of the Earth. While the mission is experimental in nature, its primary goals are to produce high quality maps and three-dimensional digital terrain models of the Earth's surface. Considerable improvement in the quality of the terrain model can be attained if information about the position and attitude of the camera is included during the adjustment of the image data. One of the primary sources of error in the Shuttle's position is due to the significant attitude maneuvers conducted during the course of the mission. Various arcs, using actual Tracking and Data Relay Satellite (TDRSS) Doppler data of STS-55, were processed to determine how effectively empirical force modeling could be used to solve for the radial, transverse, and normal components of the orbit perturbations caused by these routine maneuvers. Results are presented in terms of overlap-orbit differences in the three components. Comparisons of these differences, before and after the maneuvers are estimated, show that the quality of an orbit can be greatly enhanced with this technique, even if several maneuvers are present. Finally, a discussion is made of some of the difficulties encountered with this approach, and some ideas for future studies are presented.

Determination of Precise Satellite Orbital Position Using Multi-Band GNSS Signals

DTIC Science & Technology

2017-10-16

AFRL-AFOSR-JP-TR-2018-0002 Determination of Precise Satellite Orbital Position Using Multi -Band GNSS Signals Erry Gunawan NANYANG TECHNOLOGICAL...Position Using Multi -Band GNSS Signals 5a.  CONTRACT NUMBER 5b.  GRANT NUMBER FA2386-15-1-4041 5c.  PROGRAM ELEMENT NUMBER 61102F 6. AUTHOR(S) Erry...Grant FA2386-15-1-4041 “Determination of Precise orbital position using multi -band GNSS signals” October 13, 2017 Name of Principal Investigators

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (center) and Deputy Director Woodrow Whitlow Jr. (far left) look at the external tank door corrosion work being done on Endeavour. Next to Whitlow is Bruce Buckingham, assistant to the deputy director. Providing information, at right, are Orbiter Airframe Engineering ground area manager, and Tom Roberts, Airframe Engineering System specialist, both with United Space Alliance; and Joy Huff, with KSC Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

NASA Image and Video Library

2004-02-25

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (center) and Deputy Director Woodrow Whitlow Jr. (far left) look at the external tank door corrosion work being done on Endeavour. Next to Whitlow is Bruce Buckingham, assistant to the deputy director. Providing information, at right, are Orbiter Airframe Engineering ground area manager, and Tom Roberts, Airframe Engineering System specialist, both with United Space Alliance; and Joy Huff, with KSC Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (center) and Deputy Director Woodrow Whitlow Jr. (far left) look at the external tank door corrosion work being done on Endeavour. Next to Whitlow is Bruce Buckingham, assistant to the deputy director. Providing information, at right, are Kathy Laufenberg, Orbiter Airframe Engineering ground area manager, and Tom Roberts, Airframe Engineering System specialist, both with United Space Alliance; and Joy Huff, with Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

NASA Image and Video Library

2004-02-25

KENNEDY SPACE CENTER, FLA. - On a tour of the Orbiter Processing Facility, Center Director Jim Kennedy (center) and Deputy Director Woodrow Whitlow Jr. (far left) look at the external tank door corrosion work being done on Endeavour. Next to Whitlow is Bruce Buckingham, assistant to the deputy director. Providing information, at right, are Kathy Laufenberg, Orbiter Airframe Engineering ground area manager, and Tom Roberts, Airframe Engineering System specialist, both with United Space Alliance; and Joy Huff, with Space Shuttle Processing. Endeavour is in its Orbiter Major Modification period, which began in December 2003.

Payload/orbiter signal-processing and data-handling system evaluation

NASA Technical Reports Server (NTRS)

Teasdale, W. E.; Polydoros, A.

1980-01-01

Incompatibilities between orbiter subsystems and payload communication systems to assure that acceptable and to end system performamce will be achieved are identified. The potential incompatabilities are associated with either payloads in the cargo bay or detached payloads communicating with the orbiter via an RF link. The payload signal processing and data handling systems are assessed by investigating interface problems experienced between the inertial upper stage and the orbiter since similar problems are expected for other payloads.

Measurement of Satellite Impact Test Fragments for Modeling Orbital Debris

NASA Technical Reports Server (NTRS)

Hill, Nicole M.

2009-01-01

There are over 13,000 pieces of catalogued objects 10cm and larger in orbit around Earth [ODQN, January 2009, p12]. More than 6000 of these objects are fragments from explosions and collisions. As the earth-orbiting object count increases, debris-generating collisions in the future become a statistical inevitability. To aid in understanding this collision risk, the NASA Orbital Debris Program Office has developed computer models that calculate quantity and orbits of debris both currently in orbit and in future epochs. In order to create a reasonable computer model of the orbital debris environment, it is important to understand the mechanics of creation of debris as a result of a collision. The measurement of the physical characteristics of debris resulting from ground-based, hypervelocity impact testing aids in understanding the sizes and shapes of debris produced from potential impacts in orbit. To advance the accuracy of fragment shape/size determination, the NASA Orbital Debris Program Office recently implemented a computerized measurement system. The goal of this system is to improve knowledge and understanding of the relation between commonly used dimensions and overall shape. The technique developed involves scanning a single fragment with a hand-held laser device, measuring its size properties using a sophisticated software tool, and creating a three-dimensional computer model to demonstrate how the object might appear in orbit. This information is used to aid optical techniques in shape determination. This more automated and repeatable method provides higher accuracy in the size and shape determination of debris.

Sentry: An Automated Close Approach Monitoring System for Near-Earth Objects

NASA Astrophysics Data System (ADS)

Chamberlin, A. B.; Chesley, S. R.; Chodas, P. W.; Giorgini, J. D.; Keesey, M. S.; Wimberly, R. N.; Yeomans, D. K.

2001-11-01

In response to international concern about potential asteroid impacts on Earth, NASA's Near-Earth Object (NEO) Program Office has implemented a new system called ``Sentry'' to automatically update the orbits of all NEOs on a daily basis and compute Earth close approaches up to 100 years into the future. Results are published on our web site (http://neo.jpl.nasa.gov/) and updated orbits and ephemerides made available via the JPL Horizons ephemeris service (http://ssd.jpl.nasa.gov/horizons.html). Sentry collects new and revised astrometric observations from the Minor Planet Center (MPC) via their electronic circulars (MPECs) in near real time as well as radar and optical astrometry sent directly from observers. NEO discoveries and identifications are detected in MPECs and processed appropriately. In addition to these daily updates, Sentry synchronizes with each monthly batch of MPC astrometry and automatically updates all NEO observation files. Daily and monthly processing of NEO astrometry is managed using a queuing system which allows for manual intervention of selected NEOs without interfering with the automatic system. At the heart of Sentry is a fully automatic orbit determination program which handles outlier rejection and ensures convergence in the new solution. Updated orbital elements and their covariances are published via Horizons and our NEO web site, typically within 24 hours. A new version of Horizons, in development, will allow computation of ephemeris uncertainties using covariance data. The positions of NEOs with updated orbits are numerically integrated up to 100 years into the future and each close approach to any perturbing body in our dynamic model (all planets, Moon, Ceres, Pallas, Vesta) is recorded. Significant approaches are flagged for extended analysis including Monte Carlo studies. Results, such as minimum encounter distances and future Earth impact probabilities, are published on our NEO web site.

Taurus lightweight manned spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Chase, Kevin A.; Vandersall, Eric J.; Plotkin, Jennifer; Travisano, Jeffrey J.; Loveless, Dennis; Kaczmarek, Michael; White, Anthony G.; Est, Andy; Bulla, Gregory; Henry, Chris

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff data of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step towards larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the Space Shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster--1300 kg to a 300 km orbit. The Taurus LMS design is divided into six major design sections. The human factors system deals with the problems of life support and spacecraft cooling. The propulsion section contains the abort system, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and power generation. The thermal protection systems and spacecraft structure are contained in the structures section. The avionics section includes navigation, attitude determination, data processing, communication systems, and sensors. The mission analysis section was responsible for ground processing and spacecraft astrodynamics. The systems integration section pulled the above sections together into one spacecraft and addressed costing and reliability.

Taurus Lightweight Manned Spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Bosset, M.

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff date of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step toward larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the space shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low-cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low Earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster - 1300 kg to a 300-km orbit. The Taurus LMS design is divided into six major design sections. The Human Factors section deals with the problems of life support and spacecraft cooling. The Propulsion section contains the Abort System, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and Power Generation. The thermal protection systems and spacecraft structure are contained in the Structures section. The Avionics section includes Navigation, Attitude Determination, Data Processing, Communication systems, and Sensors. The Mission Analysis section was responsible for ground processing and spacecraft astrodynamics. The Systems Integration Section pulled the above sections together into one spacecraft, and addressed costing and reliability.

Applicability of meteor radiant determination methods depending on orbit type. I. High-eccentric orbits

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, 0.10.2 the fit is rather poor and the change of orbit unrealistic. The optimal methods with the smallest values of D for given types of orbits are shown in two series of six plots. The new method of rotation around the line of apsides we propose is very appropriate in the region of small inclinations. There is no doubt that Hasegawa's omega-adjustment method (Hasegawa, I.: 1990, Publ. Astron. Soc. Japan 42, 175) has the widest application. A comparison of the theoretical radiants with the observed radiants of seven known meteor showers is also presented.

Hot-spot model for accretion disc variability as random process. II. Mathematics of the power-spectrum break frequency

NASA Astrophysics Data System (ADS)

Pecháček, T.; Goosmann, R. W.; Karas, V.; Czerny, B.; Dovčiak, M.

2013-08-01

Context. We study some general properties of accretion disc variability in the context of stationary random processes. In particular, we are interested in mathematical constraints that can be imposed on the functional form of the Fourier power-spectrum density (PSD) that exhibits a multiply broken shape and several local maxima. Aims: We develop a methodology for determining the regions of the model parameter space that can in principle reproduce a PSD shape with a given number and position of local peaks and breaks of the PSD slope. Given the vast space of possible parameters, it is an important requirement that the method is fast in estimating the PSD shape for a given parameter set of the model. Methods: We generated and discuss the theoretical PSD profiles of a shot-noise-type random process with exponentially decaying flares. Then we determined conditions under which one, two, or more breaks or local maxima occur in the PSD. We calculated positions of these features and determined the changing slope of the model PSD. Furthermore, we considered the influence of the modulation by the orbital motion for a variability pattern assumed to result from an orbiting-spot model. Results: We suggest that our general methodology can be useful for describing non-monotonic PSD profiles (such as the trend seen, on different scales, in exemplary cases of the high-mass X-ray binary Cygnus X-1 and the narrow-line Seyfert galaxy Ark 564). We adopt a model where these power spectra are reproduced as a superposition of several Lorentzians with varying amplitudes in the X-ray-band light curve. Our general approach can help in constraining the model parameters and in determining which parts of the parameter space are accessible under various circumstances.

APOD Mission Status and Observations by VLBI

NASA Astrophysics Data System (ADS)

Tang, Geshi; Sun, Jing; Li, Xie; Liu, Shushi; Chen, Guangming; Ren, Tianpeng; Wang, Guangli

2016-12-01

On September 20, 2015, 20 satellites were successfully launched from the TaiYuan Satellite Launch Center by a Chinese CZ-6 test rocket and are, since then, operated in a circular, near-polar orbit at an altitude of 520 km. Among these satellites, a set of four CubSats, named APOD (Atmospheric density detection and Precise Orbit Determination), are intended for atmospheric density in-situ detection and derivation via precise orbit. The APOD satellites, manufactured by DFH Co., carry a number of instruments including a density detector, a dual-frequency GNSS (GPS/BD) receiver, an SLR reflector, and a VLBI S/X beacon. The APOD mission aims at detecting the atmospheric density below 520 km. The ground segment is controlled by BACC (Beijing Aerospace Control Center) including payload operation as well as science data receiving, processing, archiving, and distribution. Currently, the in-orbit test of the nano-satellites and their payloads are completed, and preliminary results show that the precision of the orbit determination is about 10 cm derived from both an overlap comparison and an SLR observation validation. The in-situ detected density calibrated by orbit-derived density demonstrates that the accuracy of atmospheric mass density is approximately 4.191×10^{-14} kgm^{-3}, about 5.5% of the measurement value. Since three space-geodetic techniques (i.e., GNSS, SLR, and VLBI) are co-located on the APOD nano-satellites, the observations can be used for combination and validation in order to detect systematic differences. Furthermore, the observations of the APOD satellites by VLBI radio telescopes can be used in an ideal fashion to link the dynamical reference frames of the satellite with the terrestrial and, most importantly, with the celestial reference frame as defined by the positions of quasars. The possibility of observing the APOD satellites by IVS VLBI radio telescopes will be analyzed, considering continental-size VLBI observing networks and the small telescopes with sufficient speed.

Application of Semi-analytical Satellite Theory orbit propagator to orbit determination for space object catalog maintenance

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.

GPS-Based Precision Orbit Determination for a New Era of Altimeter Satellites: Jason-1 and ICESat

NASA Technical Reports Server (NTRS)

Luthcke, Scott B.; Rowlands, David D.; Lemoine, Frank G.; Zelensky, Nikita P.; Williams, Teresa A.

2003-01-01

Accurate positioning of the satellite center of mass is necessary in meeting an altimeter mission's science goals. The fundamental science observation is an altimetric derived topographic height. Errors in positioning the satellite's center of mass directly impact this fundamental observation. Therefore, orbit error is a critical Component in the error budget of altimeter satellites. With the launch of the Jason-1 radar altimeter (Dec. 2001) and the ICESat laser altimeter (Jan. 2003) a new era of satellite altimetry has begun. Both missions pose several challenges for precision orbit determination (POD). The Jason-1 radial orbit accuracy goal is 1 cm, while ICESat (600 km) at a much lower altitude than Jason-1 (1300 km), has a radial orbit accuracy requirement of less than 5 cm. Fortunately, Jason-1 and ICESat POD can rely on near continuous tracking data from the dual frequency codeless BlackJack GPS receiver and Satellite Laser Ranging. Analysis of current GPS-based solution performance indicates the l-cm radial orbit accuracy goal is being met for Jason-1, while radial orbit accuracy for ICESat is well below the 54x1 mission requirement. A brief overview of the GPS precision orbit determination methodology and results for both Jason-1 and ICESat are presented.

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.

A review of GPS-based tracking techniques for TDRS orbit determination

NASA Technical Reports Server (NTRS)

Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S.-C.

1993-01-01

This article evaluates two fundamentally different approaches to the Tracking and Data Relay Satellite (TDRS) orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRS. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRS's broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and Tracking and Data Relay Satellite System satellites by ground receivers. Both strategies can be designed to meet future operational requirements for TDRS-II orbit determination.

A distributed computing approach to mission operations support. [for spacecraft

NASA Technical Reports Server (NTRS)

Larsen, R. L.

1975-01-01

Computing mission operation support includes orbit determination, attitude processing, maneuver computation, resource scheduling, etc. The large-scale third-generation distributed computer network discussed is capable of fulfilling these dynamic requirements. It is shown that distribution of resources and control leads to increased reliability, and exhibits potential for incremental growth. Through functional specialization, a distributed system may be tuned to very specific operational requirements. Fundamental to the approach is the notion of process-to-process communication, which is effected through a high-bandwidth communications network. Both resource-sharing and load-sharing may be realized in the system.

The self-calibration method for multiple systems at the CHARA Array

NASA Astrophysics Data System (ADS)

O'Brien, David

The self-calibration method, a new interferometric technique at the CHARA Array, has been used to derive orbits for several spectroscopic binaries. This method uses the wide component of a hierarchical triple system to calibrate visibility measurements of the triple's close binary system. At certain baselines and separations, the calibrator in one of these systems can be observed quasi-simultaneously with the target. Depending on the orientation of the CHARA observation baseline relative to the orientation of the wide orbit of the triple system, separated fringe packets may be observed. A sophisticated observing scheme must be put in place to ensure the existence of separated fringe packets on nights of observation. Prior to the onset of this project, the reduction of separated fringe packet data had never included the goal of deriving visibilities for both fringe packets, so new data reduction software has been written. Visibilities obtained with separated fringe packet data for the target close binary are run through both Monte Carlo simulations and grid search programs in order to determine the best-fit orbital elements of the close binary. Several targets have been observed in this fashion, and orbits have been derived for seven targets, including three new orbits. Derivation of the orbit of the close pair in a triple system allows for the calculation of the mutual inclination, which is the angle between the planes of the wide and close orbit. Knowledge of this quantity may give insight into the formation processes that create multiple star systems. INDEX WORDS: Long-baseline interferometry, Self calibration, Separated fringe packets, Triple systems, Close binaries, Multiple systems, Orbital parameters, Near-infrared interferometry

Flight Mechanics/Estimation Theory Symposium

NASA Technical Reports Server (NTRS)

Fuchs, A. J. (Editor)

1980-01-01

Methods of determining satellite orbit and attitude parameters are considered. The Goddard Trajectory Determination System, the Global Positioning System, and the Tracking and Data Relay Satellites are among the satellite navigation systems discussed. Satellite perturbation theory, orbit/attitude determination using landmark data, and star measurements are also covered.

The Outer Solar System Origin Survey full data release orbit catalog and characterization.

NASA Astrophysics Data System (ADS)

Kavelaars, J. J.; Bannister, Michele T.; Gladman, Brett; Petit, Jean-Marc; Gwyn, Stephen; Alexandersen, Mike; Chen, Ying-Tung; Volk, Kathryn; OSSOS Collaboration.

2017-10-01

The Outer Solar System Origin Survey (OSSOS) completed main data acquisition in February 2017. Here we report the release of our full orbit sample, which include 836 TNOs with high precision orbit determination and classification. We combine the OSSOS orbit sample with previously release Canada-France Ecliptic Plane Survey (CFEPS) and a precursor survey to OSSOS by Alexandersen et al. to provide a sample of over 1100 TNO orbits with high precision classified orbits and precisely determined discovery and tracking circumstances (characterization). We are releasing the full sample and characterization to the world community, along with software for conducting ‘Survey Simulations’, so that this sample of orbits can be used to test models of the formation of our outer solar system against the observed sample. Here I will present the characteristics of the data set and present a parametric model for the structure of the classical Kuiper belt.

Sentinel-2A: Orbit Modelling Improvements and their Impact on the Orbit Prediction

NASA Astrophysics Data System (ADS)

Peter, Heike; Otten, Michiel; Fernández Sánchez, Jaime; Fernández Martín, Carlos; Féménias, Pierre

2016-07-01

Sentinel-2A is the second satellite of the European Copernicus Programme. The satellite has been launched on 23rd June 2015 and it is operational since mid October 2015. This optical mission carries a GPS receiver for precise orbit determination. The Copernicus POD (Precise Orbit Determination) Service is in charge of generating precise orbital products and auxiliary files for Sentinel-2A as well as for the Sentinel-1 and -3 missions. The accuracy requirements for the Sentinel-2A orbit products are not very stringent with 3 m in 3D (3 sigma) for the near real-time (NRT) orbit and 10 m in 2D (3 sigma) for the predicted orbit. The fulfilment of the orbit accuracy requirements is normally not an issue. The Copernicus POD Service aims, however, to provide the best possible orbits for all three Sentinel missions. Therefore, a sophisticated box-wing model is generated for the Sentinel-2 satellite as it is done for the other two missions as well. Additionally, the solar wing of the satellite is rewound during eclipse, which has to be modelled accordingly. The quality of the orbit prediction is dependent on the results of the orbit estimation performed before it. The values of the last estimation of each parameter is taken for the orbit propagation, i.e. estimating ten atmospheric drag coefficients per 24h, the value of the last coefficient is used as a fix parameter for the subsequent orbit prediction. The question is whether the prediction might be stabilised by, e.g. using an average value of all ten coefficients. This paper presents the status and the quality of the Sentinel-2 orbit determination in the operational environment of the Copernicus POD Service. The impact of the orbit model improvements on the NRT and predicted orbits is studied in detail. Changes in the orbit parametrization as well as in the settings for the orbit propagation are investigated. In addition, the impact of the quality of the input GPS orbit and clock product on the Sentinel-2A orbit prediction results is checked. The results of this study do not only improve the Sentinel-2 orbit products but will also support the generation of reliable orbit predictions for the Sentinel-3 mission. The Sentinel-3 satellite is equipped with a laser retro-reflector and reliable orbit predictions are, therefore, very important to guarantee a continuous support of the satellite laser tracking stations.

Intial orbit determination results for Jason-1: towards a 1-cm orbit

NASA Technical Reports Server (NTRS)

Haines, B. J.; Haines, B.; Bertiger, W.; Desai, S.; Kuang, D.; Munson, T.; Reichert, A.; Young, L.; Willis, P.

2002-01-01

The U.S/France Jason-1 oceanographic mission is carrying state-of-the-art radiometric tracking systems (GPS and Doris) to support precise orbit determination (POD) requirements. The performance of the systems is strongly reflected in the early POD results. Results of both internal and external (e.g., satellite laser ranging) comparisons support that the 2.5 cm radial Rh4S requirement is being readily met, and provide reasons for optimism that 1 cm can be achieved. We discuss the POD strategy underlying these orbits, as well as the challenging issues that bear on the understanding and characterization of an orbit solution at the l-cm level. We also describe a system for producing science quality orbits in near real time in order to support emerging applications in operational oceanography.

Fuzzy attitude control for a nanosatellite in leo orbit

NASA Astrophysics Data System (ADS)

Calvo, Daniel; Laverón-Simavilla, Ana; Lapuerta, Victoria; Aviles, Taisir

Fuzzy logic controllers are flexible and simple, suitable for small satellites Attitude Determination and Control Subsystems (ADCS). In this work, a tailored fuzzy controller is designed for a nanosatellite and is compared with a traditional Proportional Integrative Derivative (PID) controller. Both control methodologies are compared within the same specific mission. The orbit height varies along the mission from injection at around 380 km down to a 200 km height orbit, and the mission requires pointing accuracy over the whole time. Due to both the requirements imposed by such a low orbit, and the limitations in the power available for the attitude control, a robust and efficient ADCS is required. For these reasons a fuzzy logic controller is implemented as the brain of the ADCS and its performance and efficiency are compared to a traditional PID. The fuzzy controller is designed in three separated controllers, each one acting on one of the Euler angles of the satellite in an orbital frame. The fuzzy memberships are constructed taking into account the mission requirements, the physical properties of the satellite and the expected performances. Both methodologies, fuzzy and PID, are fine-tuned using an automated procedure to grant maximum efficiency with fixed performances. Finally both methods are probed in different environments to test their characteristics. The simulations show that the fuzzy controller is much more efficient (up to 65% less power required) in single maneuvers, achieving similar, or even better, precision than the PID. The accuracy and efficiency improvement of the fuzzy controller increase with orbit height because the environmental disturbances decrease, approaching the ideal scenario. A brief mission description is depicted as well as the design process of both ADCS controllers. Finally the validation process and the results obtained during the simulations are described. Those results show that the fuzzy logic methodology is valid for small satellites' missions benefiting from a well-developed artificial intelligence theory.

On Directional Measurement Representation in Orbit Determination

DTIC Science & Technology

2016-09-13

representations. The three techniques are then compared experimentally for a geostationary and a low Earth orbit satellite using simulated data to evaluate their...Earth Orbit (LEO) and a Geostationary Earth Orbit (GEO) satellite. Section IV discusses the results from the numerical simulations and finally Section V... Geostationary Earth Orbit (GEO) satellite with the initial orbital parameters shown in Table 1. Different ground sites are used for the LEO and ahttps

KSC-98pc1166

NASA Image and Video Library

1998-09-28

The orbiter Atlantis, being towed from the Shuttle Landing Facility toward the Vehicle Assembly Building (VAB) , intersects the morning sun's rays. In the background, to the right of the VAB, are the Orbiter Processing Facility 1 and 2. Atlantis spent 10 months in Palmdale, CA, undergoing extensive inspections and modifications in the orbiter processing facility there. The modifications included several upgrades enabling it to support International Space Station missions, such as adding an external airlock for ISS docking missions and installing thinner, lighter thermal protection blankets for weight reduction which will allow it to haul heavier cargo. Atlantis will undergo preparations at KSC in Orbiter Processing Facility 2 for its planned flight in June 1999

NanoSat Constellation Mission Design

NASA Technical Reports Server (NTRS)

Concha, Marco; DeFazio, Robert

1998-01-01

The NanoSat constellation concept mission proposes simultaneous operation of multiple swarms of as many as 22 identical 10 kg spacecraft per swarm. The various orbits in a NanoSat swarm vary from 3x12 to 3x42 R(sub e) in geometry. In this report the unique flight dynamics issues of this constellation satellite mission design are addressed. Studies include orbit design, orbit determination, and error analysis. A preliminary survey determined the orbital parameters that would limit the maximum shadow condition while providing adequate ground station access for three ground stations.

Venus Gravity Handbook

NASA Technical Reports Server (NTRS)

Konopliv, Alexander S.; Sjogren, William L.

1996-01-01

This report documents the Venus gravity methods and results to date (model MGNP90LSAAP). It is called a handbook in that it contains many useful plots (such as geometry and orbit behavior) that are useful in evaluating the tracking data. We discuss the models that are used in processing the Doppler data and the estimation method for determining the gravity field. With Pioneer Venus Orbiter and Magellan tracking data, the Venus gravity field was determined complete to degree and order 90 with the use of the JPL Cray T3D Supercomputer. The gravity field shows unprecedented high correlation with topography and resolution of features to the 2OOkm resolution. In the procedure for solving the gravity field, other information is gained as well, and, for example, we discuss results for the Venus ephemeris, Love number, pole orientation of Venus, and atmospheric densities. Of significance is the Love number solution which indicates a liquid core for Venus. The ephemeris of Venus is determined to an accuracy of 0.02 mm/s (tens of meters in position), and the rotation period to 243.0194 +/- 0.0002 days.

Kalman filtering applied to real-time monitoring of apogee maneuvers

NASA Technical Reports Server (NTRS)

Deboer, Frederic; Barbier, Christian

1993-01-01

Part of the Space Mathematics Division in CNES, the Flight Dynamics Center provides attitude and orbit determinations and maneuvers during the Launch and Early Operation Phase (LEOP) of geostationary satellites. Orbit determination is based on a Kalman filter method; when the 2 GHz CNES/NASA network is used, Doppler measurements are available and allow orbit determination during the apogee maneuvers. This method was used for TELE-X and TDF 2 LEOP (3-axis controlled satellites) and also for TELECOM 2 and HISPASAT (spun satellites): it enables us to follow the evolution of the maneuver and gives out a quite accurate estimation of the reached orbit. In this paper, we briefly describe the dynamic models of the orbit evolution in both cases, '3-axis' and 'inertial' thrust. Then, we present the results obtained for each case. Afterwards, we present some cases to show the robustness of the filter.

Simulation of interference between Earth stations and Earth-orbiting satellites

NASA Technical Reports Server (NTRS)

Bishop, D. F.

1994-01-01

It is often desirable to determine the potential for radio frequency interference between earth stations and orbiting spacecraft. This information can be used to select frequencies for radio systems to avoid interference or it can be used to determine if coordination between radio systems is necessary. A model is developed that will determine the statistics of interference between earth stations and elliptical orbiting spacecraft. The model uses orbital dynamics, detailed antenna patterns, and spectral characteristics to obtain accurate levels of interference at the victim receiver. The model is programmed into a computer simulation to obtain long-term statistics of interference. Two specific examples are shown to demonstrate the model. The first example is a simulation of interference from a fixed-satellite earth station to an orbiting scatterometer receiver. The second example is a simulation of interference from earth-exploration satellites to a deep-space earth station.

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.

Orbit determination based on meteor observations using numerical integration of equations of motion

NASA Astrophysics Data System (ADS)

Dmitriev, V.; Lupovka, V.; Gritsevich, M.

2014-07-01

We review the definitions and approaches to orbital-characteristics analysis applied to photographic or video ground-based observations of meteors. A number of camera networks dedicated to meteors registration were established all over the word, including USA, Canada, Central Europe, Australia, Spain, Finland and Poland. Many of these networks are currently operational. The meteor observations are conducted from different locations hosting the network stations. Each station is equipped with at least one camera for continuous monitoring of the firmament (except possible weather restrictions). For registered multi-station meteors, it is possible to accurately determine the direction and absolute value for the meteor velocity and thus obtain the topocentric radiant. Based on topocentric radiant one further determines the heliocentric meteor orbit. We aim to reduce total uncertainty in our orbit-determination technique, keeping it even less than the accuracy of observations. The additional corrections for the zenith attraction are widely in use and are implemented, for example, here [1]. We propose a technique for meteor-orbit determination with higher accuracy. We transform the topocentric radiant in inertial (J2000) coordinate system using the model recommended by IAU [2]. The main difference if compared to the existing orbit-determination techniques is integration of ordinary differential equations of motion instead of addition correction in visible velocity for zenith attraction. The attraction of the central body (the Sun), the perturbations by Earth, Moon and other planets of the Solar System, the Earth's flattening (important in the initial moment of integration, i.e. at the moment when a meteoroid enters the atmosphere), atmospheric drag may be optionally included in the equations. In addition, reverse integration of the same equations can be performed to analyze orbital evolution preceding to meteoroid's collision with Earth. To demonstrate the developed technique, we provide calculated orbits for several cases, including well-known meteorite-producing fireballs. A comparison of our estimates with previously published ones is also provided.

TerraSAR-X precise orbit determination with real-time GPS ephemerides

NASA Astrophysics Data System (ADS)

Wermuth, Martin; Hauschild, Andre; Montenbruck, Oliver; Kahle, Ralph

TerraSAR-X is a German Synthetic Aperture Radar (SAR) satellite, which was launched in June 2007 from Baikonour. Its task is to acquire radar images of the Earth's surface. In order to locate the radar data takes precisely, the satellite is equipped with a high-quality dual-frequency GPS receiver -the Integrated Geodetic and Occultation Receiver (IGOR) provided by the GeoForschungsZentrum Potsdam (GFZ). Using GPS observations from the IGOR instrument in a reduced dynamic precise orbit determination (POD), the German Space Operations Center (DLR/GSOC) is computing rapid and science orbit products on a routine basis. The rapid orbit products arrive with a latency of about one hour after data reception with an accuracy of 10-20 cm. Science orbit products are computed with a latency of five days achieving an accuracy of about 5cm (3D-RMS). For active and future Earth observation missions, the availability of near real-time precise orbit information is becoming more and more important. Other applications of near real-time orbit products include the processing of GNSS radio occulation measurements for atmospheric sounding as well as altimeter measurements of ocean surface heights, which are nowadays employed in global weather and ocean circulation models with short latencies. For example after natural disasters it is necessary to evaluate the damage by satellite images as soon as possible. The latency and quality of POD results is mainly driven by the availability of precise GPS ephemerides. In order to have high-quality GPS ephemerides available at real-time, GSOC has developed the real-time clock estimation system RETICLE. The system receives NTRIP-data streams with GNSS observations from the global tracking network of IGS in real-time. Using the known station position, RETICLE estimates precise GPS satellite clock offsets and drifts based on the most recent available IGU predicted orbits. The clock offset estimates have an accuracy of better than 0.3 ns and are globally valid. The latency of the estimated clocks is approximately 7 seconds. Another limiting factor is the frequency of satellite downlinks and the latency of the data transfer from the ground station to the computation center. Therefore a near real-time scenario is examined in which the satellite has about one ground station contact per orbit or respectively one contact in 90 minutes. The results of the near real-time POD are evaluated in an internal consistency check and compared against the science orbit solution and laser ranging observations.

Stability of outer planetary orbits around binary stars - A comparison of Hill's and Laplace's stability criteria

NASA Technical Reports Server (NTRS)

Kubala, A.; Black, D.; Szebehely, V.

1993-01-01

A comparison is made between the stability criteria of Hill and that of Laplace to determine the stability of outer planetary orbits encircling binary stars. The restricted, analytically determined results of Hill's method by Szebehely and coworkers and the general, numerically integrated results of Laplace's method by Graziani and Black (1981) are compared for varying values of the mass parameter mu. For mu = 0 to 0.15, the closest orbit (lower limit of radius) an outer planet in a binary system can have and still remain stable is determined by Hill's stability criterion. For mu greater than 0.15, the critical radius is determined by Laplace's stability criterion. It appears that the Graziani-Black stability criterion describes the critical orbit within a few percent for all values of mu.

Precision analysis of autonomous orbit determination using star sensor for Beidou MEO satellite

NASA Astrophysics Data System (ADS)

Shang, Lin; Chang, Jiachao; Zhang, Jun; Li, Guotong

2018-04-01

This paper focuses on the autonomous orbit determination accuracy of Beidou MEO satellite using the onboard observations of the star sensors and infrared horizon sensor. A polynomial fitting method is proposed to calibrate the periodic error in the observation of the infrared horizon sensor, which will greatly influence the accuracy of autonomous orbit determination. Test results show that the periodic error can be eliminated using the polynomial fitting method. The User Range Error (URE) of Beidou MEO satellite is less than 2 km using the observations of the star sensors and infrared horizon sensor for autonomous orbit determination. The error of the Right Ascension of Ascending Node (RAAN) is less than 60 μrad and the observations of star sensors can be used as a spatial basis for Beidou MEO navigation constellation.

Fuel-Optimal Altitude Maintenance of Low-Earth-Orbit Spacecrafts by Combined Direct/Indirect Optimization

NASA Astrophysics Data System (ADS)

Kim, Kyung-Ha; Park, Chandeok; Park, Sang-Young

2015-12-01

This work presents fuel-optimal altitude maintenance of Low-Earth-Orbit (LEO) spacecrafts experiencing non-negligible air drag and J2 perturbation. A pseudospectral (direct) method is first applied to roughly estimate an optimal fuel consumption strategy, which is employed as an initial guess to precisely determine itself. Based on the physical specifications of KOrea Multi-Purpose SATellite-2 (KOMPSAT-2), a Korean artificial satellite, numerical simulations show that a satellite ascends with full thrust at the early stage of the maneuver period and then descends with null thrust. While the thrust profile is presumably bang-off, it is difficult to precisely determine the switching time by using a pseudospectral method only. This is expected, since the optimal switching epoch does not coincide with one of the collocation points prescribed by the pseudospectral method, in general. As an attempt to precisely determine the switching time and the associated optimal thrust history, a shooting (indirect) method is then employed with the initial guess being obtained through the pseudospectral method. This hybrid process allows the determination of the optimal fuel consumption for LEO spacecrafts and their thrust profiles efficiently and precisely.

KENNEDY SPACE CENTER, FLA. - STS-114 Mission Specialist Wendy Lawrence autographs the sign presented to workers in the Orbiter Processing Facility. Lawrence is a new addition to the crew. The STS-114 crew is at KSC to take part in crew equipment and orbiter familiarization.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - STS-114 Mission Specialist Wendy Lawrence autographs the sign presented to workers in the Orbiter Processing Facility. Lawrence is a new addition to the crew. The STS-114 crew is at KSC to take part in crew equipment and orbiter familiarization.

Impact of Multi-GNSS Observations on Precise Orbit Determination and Precise Point Positioning Solutions

NASA Astrophysics Data System (ADS)

Amiri, N.; Bertiger, W. I.; Lu, W.; Miller, M. A.; David, M. W.; Ries, P.; Romans, L.; Sibois, A. E.; Sibthorpe, A.; Sakumura, C.

2017-12-01

Impact of Multi-GNSS Observations on Precise Orbit Determination and Precise Point Positioning Solutions Authors: Nikta Amiri, Willy Bertiger, Wenwen Lu, Mark Miller, David Murphy, Paul Ries, Larry Romans, Carly Sakumura, Aurore Sibois, Anthony Sibthorpe All at the Jet Propulsion Laboratory, California Institute of Technology Multiple Global Navigation Satellite Systems (GNSS) are now in various stages of completion. The four current constellations (GPS, GLONASS, BeiDou, Galileo) comprise more than 80 satellites as of July 2017, with 120 satellites expected to be available when all four constellations become fully operational. We investigate the impact of simultaneous observations to these four constellations on global network precise orbit determination (POD) solutions, and compare them to available sets of orbit and clock products submitted to the Multi-GNSS Experiment (MGEX). Using JPL's GipsyX software, we generate orbit and clock products for the four constellations. The resulting solutions are evaluated based on a number of metrics including day-to-day internal and external orbit and/or clock overlaps and estimated constellation biases. Additionally, we examine estimated station positions obtained from precise point positioning (PPP) solutions by comparing results generated from multi-GNSS and GPS-only orbit and clock products.

Navigation Algorithms for the SeaWiFS Mission

NASA Technical Reports Server (NTRS)

Hooker, Stanford B. (Editor); Firestone, Elaine R. (Editor); Patt, Frederick S.; McClain, Charles R. (Technical Monitor)

2002-01-01

The navigation algorithms for the Sea-viewing Wide Field-of-view Sensor (SeaWiFS) were designed to meet the requirement of 1-pixel accuracy-a standard deviation (sigma) of 2. The objective has been to extract the best possible accuracy from the spacecraft telemetry and avoid the need for costly manual renavigation or geometric rectification. The requirement is addressed by postprocessing of both the Global Positioning System (GPS) receiver and Attitude Control System (ACS) data in the spacecraft telemetry stream. The navigation algorithms described are separated into four areas: orbit processing, attitude sensor processing, attitude determination, and final navigation processing. There has been substantial modification during the mission of the attitude determination and attitude sensor processing algorithms. For the former, the basic approach was completely changed during the first year of the mission, from a single-frame deterministic method to a Kalman smoother. This was done for several reasons: a) to improve the overall accuracy of the attitude determination, particularly near the sub-solar point; b) to reduce discontinuities; c) to support the single-ACS-string spacecraft operation that was started after the first mission year, which causes gaps in attitude sensor coverage; and d) to handle data quality problems (which became evident after launch) in the direct-broadcast data. The changes to the attitude sensor processing algorithms primarily involved the development of a model for the Earth horizon height, also needed for single-string operation; the incorporation of improved sensor calibration data; and improved data quality checking and smoothing to handle the data quality issues. The attitude sensor alignments have also been revised multiple times, generally in conjunction with the other changes. The orbit and final navigation processing algorithms have remained largely unchanged during the mission, aside from refinements to data quality checking. Although further improvements are certainly possible, future evolution of the algorithms is expected to be limited to refinements of the methods presented here, and no substantial changes are anticipated.

Abort Options for Human Missions to Earth-Moon Halo Orbits

NASA Technical Reports Server (NTRS)

Jesick, Mark C.

2013-01-01

Abort trajectories are optimized for human halo orbit missions about the translunar libration point (L2), with an emphasis on the use of free return trajectories. Optimal transfers from outbound free returns to L2 halo orbits are numerically optimized in the four-body ephemeris model. Circumlunar free returns are used for direct transfers, and cislunar free returns are used in combination with lunar gravity assists to reduce propulsive requirements. Trends in orbit insertion cost and flight time are documented across the southern L2 halo family as a function of halo orbit position and free return flight time. It is determined that the maximum amplitude southern halo incurs the lowest orbit insertion cost for direct transfers but the maximum cost for lunar gravity assist transfers. The minimum amplitude halo is the most expensive destination for direct transfers but the least expensive for lunar gravity assist transfers. The on-orbit abort costs for three halos are computed as a function of abort time and return time. Finally, an architecture analysis is performed to determine launch and on-orbit vehicle requirements for halo orbit missions.

Improving multi-GNSS ultra-rapid orbit determination for real-time precise point positioning

NASA Astrophysics Data System (ADS)

Li, Xingxing; Chen, Xinghan; Ge, Maorong; Schuh, Harald

2018-03-01

Currently, with the rapid development of multi-constellation Global Navigation Satellite Systems (GNSS), the real-time positioning and navigation are undergoing dramatic changes with potential for a better performance. To provide more precise and reliable ultra-rapid orbits is critical for multi-GNSS real-time positioning, especially for the three merging constellations Beidou, Galileo and QZSS which are still under construction. In this contribution, we present a five-system precise orbit determination (POD) strategy to fully exploit the GPS + GLONASS + BDS + Galileo + QZSS observations from CDDIS + IGN + BKG archives for the realization of hourly five-constellation ultra-rapid orbit update. After adopting the optimized 2-day POD solution (updated every hour), the predicted orbit accuracy can be obviously improved for all the five satellite systems in comparison to the conventional 1-day POD solution (updated every 3 h). The orbit accuracy for the BDS IGSO satellites can be improved by about 80, 45 and 50% in the radial, cross and along directions, respectively, while the corresponding accuracy improvement for the BDS MEO satellites reaches about 50, 20 and 50% in the three directions, respectively. Furthermore, the multi-GNSS real-time precise point positioning (PPP) ambiguity resolution has been performed by using the improved precise satellite orbits. Numerous results indicate that combined GPS + BDS + GLONASS + Galileo (GCRE) kinematic PPP ambiguity resolution (AR) solutions can achieve the shortest time to first fix (TTFF) and highest positioning accuracy in all coordinate components. With the addition of the BDS, GLONASS and Galileo observations to the GPS-only processing, the GCRE PPP AR solution achieves the shortest average TTFF of 11 min with 7{°} cutoff elevation, while the TTFF of GPS-only, GR, GE and GC PPP AR solution is 28, 15, 20 and 17 min, respectively. As the cutoff elevation increases, the reliability and accuracy of GPS-only PPP AR solutions decrease dramatically, but there is no evident decrease for the accuracy of GCRE fixed solutions which can still achieve an accuracy of a few centimeters in the east and north components.

Thermal models applicable for visual and infrared studies of orbital debris

NASA Technical Reports Server (NTRS)

Lebofsky, Larry A.; Vilas, Faith

1990-01-01

Over the past decade, thermal models have been developed for the determination of asteroid diameters and albedos. As a first step to understanding the size/frequency distribution of the debris population in earth orbit, these thermal models have been modified to determine the sizes of orbiting debris. When possible, the model results have been compared to spherical satellites of known diameter.

KSC-04pd0950

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is lowered toward the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KSC-04pd0948

NASA Image and Video Library

2004-04-21

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the left-hand Orbital Maneuvering System (OMS) pod is maneuvered toward the engine interfaces on the orbiter Discovery for installation. The Orbital Maneuvering System provides the thrust for orbit insertion, orbit circularization, orbit transfer, rendezvous, deorbit, abort to orbit and abort once around. It can provide up to 1,000 pounds of propellant to the aft reaction control system. Each pod contains one OMS engine and the hardware needed to pressurize, store and distribute the propellants to perform the velocity maneuvers.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Joel Smith prepares an area on the orbiter Discovery for blanket installation. The blankets are part of the Orbiter Thermal Protection System, thermal shields to protect against temperatures as high as 3,000° Fahrenheit, which are produced during descent for landing. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

NASA Image and Video Library

2003-12-09

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Joel Smith prepares an area on the orbiter Discovery for blanket installation. The blankets are part of the Orbiter Thermal Protection System, thermal shields to protect against temperatures as high as 3,000° Fahrenheit, which are produced during descent for landing. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Nadine Phillips prepares an area on the orbiter Discovery for blanket installation. The blankets are part of the Orbiter Thermal Protection System, thermal shields to protect against temperatures as high as 3,000° Fahrenheit, which are produced during descent for landing. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

NASA Image and Video Library

2003-12-09

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, KSC employee Nadine Phillips prepares an area on the orbiter Discovery for blanket installation. The blankets are part of the Orbiter Thermal Protection System, thermal shields to protect against temperatures as high as 3,000° Fahrenheit, which are produced during descent for landing. Discovery is scheduled to fly on mission STS-121 to the International Space Station.

Orbit computation of the TELECOM-2D satellite with a Genetic Algorithm

NASA Astrophysics Data System (ADS)

Deleflie, Florent; Coulot, David; Vienne, Alain; Decosta, Romain; Richard, Pascal; Lasri, Mohammed Amjad

2014-07-01

In order to test a preliminary orbit determination method, we fit an orbit of the geostationary satellite TELECOM-2D, as if we did not know any a priori information on its trajectory. The method is based on a genetic algorithm coupled to an analytical propagator of the trajectory, that is used over a couple of days, and that uses a whole set of altazimutal data that are acquired by the tracking network made up of the two TAROT telescopes. The adjusted orbit is then compared to a numerical reference. The method is described, and the results are analyzed, as a step towards an operational method of preliminary orbit determination for uncatalogued objects.

GEODYN system support program, volume 4. [computer program for trajectory analysis of artificial satellites

NASA Technical Reports Server (NTRS)

Mullins, N. E.

1972-01-01

The GEODYN Orbit Determination and Geodetic Parameter Estimation System consists of a set of computer programs designed to determine and analyze definitive satellite orbits and their associated geodetic and measurement parameters. This manual describes the Support Programs used by the GEODYN System. The mathematics and programming descriptions are detailed. The operational procedures of each program are presented. GEODYN ancillary analysis programs may be grouped into three different categories: (1) orbit comparison - DELTA (2) data analysis using reference orbits - GEORGE, and (3) pass geometry computations - GROUNDTRACK. All of the above three programs use one or more tapes written by the GEODYN program in either a data reduction or orbit generator run.

Rosetta mission operations for landing

NASA Astrophysics Data System (ADS)

Accomazzo, Andrea; Lodiot, Sylvain; Companys, Vicente

2016-08-01

The International Rosetta Mission of the European Space Agency (ESA) was launched on 2nd March 2004 on its 10 year journey to comet Churyumov-Gerasimenko and has reached it early August 2014. The main mission objectives were to perform close observations of the comet nucleus throughout its orbit around the Sun and deliver the lander Philae to its surface. This paper describers the activities at mission operations level that allowed the landing of Philae. The landing preparation phase was mainly characterised by the definition of the landing selection process, to which several parties contributed, and by the definition of the strategy for comet characterisation, the orbital strategy for lander delivery, and the definition and validation of the operations timeline. The definition of the landing site selection process involved almost all components of the mission team; Rosetta has been the first, and so far only mission, that could not rely on data collected by previous missions for the landing site selection. This forced the teams to include an intensive observation campaign as a mandatory part of the process; several science teams actively contributed to this campaign thus making results from science observations part of the mandatory operational products. The time allocated to the comet characterisation phase was in the order of a few weeks and all the processes, tools, and interfaces required an extensive planning an validation. Being the descent of Philae purely ballistic, the main driver for the orbital strategy was the capability to accurately control the position and velocity of Rosetta at Philae's separation. The resulting operations timeline had to merge this need of frequent orbit determination and control with the complexity of the ground segment and the inherent risk of problems when doing critical activities in short times. This paper describes the contribution of the Mission Control Centre (MOC) at the European Space Operations Centre (ESOC) to this mission phase and the lessons learned that can be derived from this experience.

Effects of solar radiation on the orbits of small particles

NASA Technical Reports Server (NTRS)

Lyttleton, R. A.

1976-01-01

A modification of the Robertson (1937) equations of particle motion in the presence of solar radiation is developed which allows for partial reflection of sunlight as a result of rapid and varying particle rotations caused by interaction with the solar wind. The coefficients and forces in earlier forms of the equations are compared with those in the present equations, and secular rates of change of particle orbital elements are determined. Orbital dimensions are calculated in terms of time, probable sizes and densities of meteoric and cometary particles are estimated, and times of infall to the sun are computed for a particle moving in an almost circular orbit and a particle moving in an elliptical orbit of high eccentricity. Changes in orbital elements are also determined for particles from a long-period sun-grazing comet. The results show that the time of infall to the sun from a highly eccentric orbit is substantially shorter than from a circular orbit with a radius equal to the mean distance in the eccentric orbit. The possibility is considered that the free orbital kinetic energy of particles drawn into the sun may be the energy source for the solar corona.

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 Modernized Approach to Meet Diversified Earth Observing System (EOS) AM-1 Mission Requirements

NASA Technical Reports Server (NTRS)

Newman, Lauri Kraft; Hametz, Mark E.; Conway, Darrel J.

1998-01-01

From a flight dynamics perspective, the EOS AM-1 mission design and maneuver operations present a number of interesting challenges. The mission design itself is relatively complex for a low Earth mission, requiring a frozen, Sun-synchronous, polar orbit with a repeating ground track. Beyond the need to design an orbit that meets these requirements, the recent focus on low-cost, "lights out" operations has encouraged a shift to more automated ground support. Flight dynamics activities previously performed in special facilities created solely for that purpose and staffed by personnel with years of design experience are now being shifted to the mission operations centers (MOCs) staffed by flight operations team (FOT) operators. These operators' responsibilities include flight dynamics as a small subset of their work; therefore, FOT personnel often do not have the experience to make critical maneuver design decisions. Thus, streamlining the analysis and planning work required for such a complicated orbit design and preparing FOT personnel to take on the routine operation of such a spacecraft both necessitated increasing the automation level of the flight dynamics functionality. The FreeFlyer(trademark) software developed by AI Solutions provides a means to achieve both of these goals. The graphic interface enables users to interactively perform analyses that previously required many parametric studies and much data reduction to achieve the same result. In addition, the fuzzy logic engine .enables the simultaneous evaluation of multiple conflicting constraints, removing the analyst from the loop and allowing the FOT to perform more of the operations without much background in orbit design. Modernized techniques were implemented for EOS AM-1 flight dynamics support in several areas, including launch window determination, orbit maintenance maneuver control strategies, and maneuver design and calibration automation. The benefits of implementing these techniques include increased fuel available for on-orbit maneuvering, a simplified orbit maintenance process to minimize science data downtime, and an automated routine maneuver planning process. This paper provides an examination of the modernized techniques implemented for EOS AM-1 to achieve these benefits.

A modernized approach to meet diversified earth observing system (EOS) AM-1 mission requirements

NASA Technical Reports Server (NTRS)

Newman, Lauri Kraft; Hametz, Mark E.; Conway, Darrel J.

1998-01-01

From a flight dynamics perspective, the EOS AM-1 mission design and maneuver operations present a number of interesting challenges. The mission design itself is relatively complex for a low Earth mission, requiring a frozen, Sun-synchronous, polar orbit with a repeating ground track. Beyond the need to design an orbit that meets these requirements, the recent focus on low-cost, 'lights out' operations has encouraged a shift to more automated ground support. Flight dynamics activities previously performed in special facilities created solely for that purpose and staffed by personnel with years of design experience are now being shifted to the mission operations centers (MOCs) staffed by flight operations team (FOT) operators. These operators' responsibilities include flight dynamics as a small subset of their work; therefore, FOT personnel often do not have the experience to make critical maneuver design decisions. Thus, streamlining the analysis and planning work required for such a complicated orbit design and preparing FOT personnel to take on the routine operation of such a spacecraft both necessitated increasing the automation level of the flight dynamics functionality. The FreeFlyer(TM) software developed by AI Solutions provides a means to achieve both of these goals. The graphic interface enables users to interactively perform analyses that previously required many parametric studies and much data reduction to achieve the same result In addition, the fuzzy logic engine enables the simultaneous evaluation of multiple conflicting constraints, removing the analyst from the loop and allowing the FOT to perform more of the operations without much background in orbit design. Modernized techniques were implemented for EOS AM-1 flight dynamics support in several areas, including launch window determination, orbit maintenance maneuver control strategies, and maneuver design and calibration automation. The benefits of implementing these techniques include increased fuel available for on-orbit maneuvering, a simplified orbit maintenance process to minimize science data downtime, and an automated routine maneuver planning process. This paper provides an examination of the modernized techniques implemented for EOS AM-1 to achieve these benefits.

Methods for extracting aerodynamic accelerations from Orbiter High Resolution Accelerometer Package flight data

NASA Technical Reports Server (NTRS)

Thompson, J. M.; Russell, J. W.; Blanchard, R. C.

1987-01-01

This report presents a process for extracting the aerodynamic accelerations of the Shuttle Orbiter Vehicle from the High Resolution Accelerometer Package (HiRAP) flight data during reentry. The methods for obtaining low-level aerodynamic accelerations, principally in the rarefied flow regime, are applied to 10 Orbiter flights. The extraction process is presented using data obtained from Space Transportation System Flight 32 (Mission 61-C) as a typical example. This process involves correcting the HiRAP measurements for the effects of temperature bias and instrument offset from the Orbiter center of gravity, and removing acceleration data during times they are affected by thruster firings. The corrected data are then made continuous and smooth and are further enhanced by refining the temperature bias correction and removing effects of the auxiliary power unit actuation. The resulting data are the current best estimate of the Orbiter aerodynamic accelerations during reentry and will be used for further analyses of the Orbiter aerodynamics and the upper atmosphere characteristics.

Shuttle on-orbit rendezvous targeting: Circular orbits

NASA Technical Reports Server (NTRS)

Bentley, E. L.

1972-01-01

The strategy and logic used in a space shuttle on-orbit rendezvous targeting program are described. The program generates ascent targeting conditions for boost to insertion into an intermediate parking orbit, and generates on-orbit targeting and timeline bases for each maneuver to effect rendezvous with a space station. Time of launch is determined so as to eliminate any plane change, and all work was performed for a near-circular space station orbit.

Integrated orbital servicing study for low-cost payload programs. Volume 2: Technical and cost analysis

NASA Technical Reports Server (NTRS)

Cody, E. R.; Deats, C. L.; Derocher, W. L., Jr.; Kyrias, G. M.; Snodgrass, M. R.; Sosnay, R. D.; Spencer, R. A.; Wudell, A. E.

1975-01-01

Orbital maintenance concepts were examined in an effort to determine a cost effective orbital maintenance system compatible with the space transportation system. An on-orbit servicer maintenance system is recommended as the most cost effective system. A pivoting arm on-orbit servicer was selected and a preliminary design was prepared. It is indicated that orbital maintenance does not have any significant impact on the space transportation system.

STK/Lifetime as a Replacement for Heritage Orbital Lifetime Software

NASA Technical Reports Server (NTRS)

Dove, Edwin

2004-01-01

The Flight Dynamics Analysis Branch (FDAB) of NASNGSFC is tasked with determining the orbital lifetime of several developmental and operational satellites, which include the Hubble Space Telescope. A DOS based program developed by the FDAB many years ago, called PC Lifetime, is used to determine a satellite s lifetime and could soon be in need of a replacement. STK s Lifetime Object Tool is a possible candidate. Due to the reduced support of the PC Lifetime program, and the growing incompatibility of older programs with new operating systems, a comparative analysis was done to determine if STWLifetime could meet the stringent requirements that were laid before it. The use of highly accurate numerical propagators such as STK s High Precision Orbit Propagator ( OP) and the Goddard Trajectory Determination System (GTDS) provided a basis on which to compare STWLifetime s results. Several test cases were run, but the main four test cases would determine whether or not STWLifetime could be PC- Lifetime s replacement. These four cases include a geotransfer orbit, two circular LEOS, and a Poiar LEO. Following rigorous testmg procedures, a conclusion will be determined. STK has proved to be a versatile program on many satellite missions and the FDAB has high hopes that it can pass FDAB s requirements for orbital lifetime prediction.

Nontraditional method for determining unperturbed orbits of unknown space objects using incomplete optical observational data

NASA Astrophysics Data System (ADS)

Perov, N. I.

1985-02-01

A physical-geometrical method for computing the orbits of earth satellites on the basis of an inadequate number of angular observations (N3) was developed. Specifically, a new method has been developed for calculating the elements of Keplerian orbits of unidentified artificial satellites using two angular observations (alpha sub k, S sub k, k = 1). The first section gives procedures for determining the topocentric distance to AES on the basis of one optical observation. This is followed by description of a very simple method for determining unperturbed orbits using two satellite position vectors and a time interval which is applicable even in the case of antiparallel AED position vectors, a method designated the R sub 2 iterations method.

Separation in 5 Msun Binaries

NASA Astrophysics Data System (ADS)

Evans, Nancy R.; Bond, H. E.; Schaefer, G.; Mason, B. D.; Karovska, M.; Tingle, E.

2013-01-01

Cepheids (5 Msun stars) provide an excellent sample for determining the binary properties of fairly massive stars. International Ultraviolet Explorer (IUE) observations of Cepheids brighter than 8th magnitude resulted in a list of ALL companions more massive than 2.0 Msun uniformly sensitive to all separations. Hubble Space Telescope Wide Field Camera 3 (WFC3) has resolved three of these binaries (Eta Aql, S Nor, and V659 Cen). Combining these separations with orbital data in the literature, we derive an unbiased distribution of binary separations for a sample of 18 Cepheids, and also a distribution of mass ratios. The distribution of orbital periods shows that the 5 Msun binaries prefer shorter periods than 1 Msun stars, reflecting differences in star formation processes.

Comet C/2017 K2 (PANSTARRS): Dynamically Old or New?

NASA Astrophysics Data System (ADS)

de la Fuente Marcos, Raúl; de la Fuente Marcos, Carlos

2018-04-01

At discovery time, C/2017 K2 (PANSTARRS) was the second most distant inbound active comet ever observed. It has been argued that this object is in the process of crossing the inner Solar System for the first time, but other authors have concluded that it is dynamically old. We have performed full N-body simulations for 3 Myr into the past using the latest public orbit determination for this object and most of them, 67%, are consistent with a bound and dynamically old Oort cloud comet, but about 29% of the studied orbits are compatible with an interstellar origin. Our independent calculations strongly suggest that C/2017 K2 is not a dynamically new Oort cloud comet.

Study on Synergistic Mechanism of Inhibitor Mixture Based on Electron Transfer Behavior

PubMed Central

Han, Peng; He, Yang; Chen, Changfeng; Yu, Haobo; Liu, Feng; Yang, Hong; Ma, Yue; Zheng, Yanjun

2016-01-01

Mixing is an important method to improve the performance of surfactants due to their synergistic effect. The changes in bonding interaction and adsorption structure of IM and OP molecules before and after co-adsorbed on Fe(001) surface is calculated by DFTB+ method. It is found that mixture enable the inhibitor molecules with higher EHOMO donate more electrons while the inhibitor molecules with lower ELUMO accept more electrons, which strengthens the bonding interaction of both inhibitor agent and inhibitor additive with metal surface. Meanwhile, water molecules in the compact layer of double electric layer are repulsed and the charge transfer resistance during the corrosion process increases. Accordingly, the correlation between the frontier orbital (EHOMO and ELUMO of inhibitor molecules and the Fermi level of metal) and inhibition efficiency is determined. Finally, we propose a frontier orbital matching principle for the synergistic effect of inhibitors, which is verified by electrochemical experiments. This frontier orbital matching principle provides an effective quantum chemistry calculation method for the optimal selection of inhibitor mixture. PMID:27671332

Vibronic coupling effect on the electron transport through molecules

NASA Astrophysics Data System (ADS)

Tsukada, Masaru; Mitsutake, Kunihiro

2007-03-01

Electron transport through molecular bridges or molecular layers connected to nano-electrodes is determined by the combination of coherent and dissipative processes, controlled by the electron-vibron coupling, transfer integrals between the molecular orbitals, applied electric field and temperature. We propose a novel theoretical approach, which combines ab initio molecular orbital method with analytical many-boson model. As a case study, the long chain model of the thiophene oligomer is solved by a variation approach. Mixed states of moderately extended molecular orbital states mediated and localised by dress of vibron cloud are found as eigen-states. All the excited states accompanied by multiple quanta of vibration can be solved, and the overall carrier transport properties including the conductance, mobility, dissipation spectra are analyzed by solving the master equation with the transition rates estimated by the golden rule. We clarify obtained in a uniform systematic way, how the transport mode changes from a dominantly coherent transport to the dissipative hopping transport.

Mars Aeronomy Explorer (MAX): Study Employing Distributed Micro-Spacecraft

NASA Technical Reports Server (NTRS)

Shotwell, Robert F.; Gray, Andrew A.; Illsley, Peter M.; Johnson, M.; Sherwood, Robert L.; Vozoff, M.; Ziemer, John K.

2005-01-01

An overview of a Mars Aeronomy Explorer (MAX) mission design study performed at NASA's Jet Propulsion Laboratory is presented herein. The mission design consists of ten micro-spacecraft orbiters launched on a Delta IV to Mars polar orbit to determine the spatial, diurnal and seasonal variation of the constituents of the Martian upper atmosphere and ionosphere over the course of one Martian year. The spacecraft are designed to allow penetration of the upper atmosphere to at least 90 km. This property coupled with orbit precession will yield knowledge of the nature of the solar wind interaction with Mars, the influence of the Mars crustal magnetic field on ionospheric processes, and the measurement of present thermal and nonthermal escape rates of atmospheric constituents. The mission design incorporates alternative design paradigms that are more appropriate for-and in some cases motivate-distributed micro-spacecraft. These design paradigms are not defined by a simple set of rules, but rather a way of thinking about the function of instruments, mission reliability/risk, and cost in a systemic framework.

Independent Orbiter Assessment (IOA): Analysis of the atmospheric revitalization pressure control subsystem

NASA Technical Reports Server (NTRS)

Saiidi, M. J.; Duffy, R. E.; Mclaughlin, T. D.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis/Critical Items List (FMEA/CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results corresponding to the Orbiter Atmospheric Revitalization and Pressure Control Subsystem (ARPCS) are documented. The ARPCS hardware was categorized into the following subdivisions: (1) Atmospheric Make-up and Control (including the Auxiliary Oxygen Assembly, Oxygen Assembly, and Nitrogen Assembly); and (2) Atmospheric Vent and Control (including the Positive Relief Vent Assembly, Negative Relief Vent Assembly, and Cabin Vent Assembly). The IOA analysis process utilized available ARPCS hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Independent Orbiter Assessment (IOA): Analysis of the mechanical actuation subsystem

NASA Technical Reports Server (NTRS)

Bacher, J. L.; Montgomery, A. D.; Bradway, M. W.; Slaughter, W. T.

1987-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Mechanical Actuation System (MAS) hardware. Specifically, the MAS hardware consists of the following components: Air Data Probe (ADP); Elevon Seal Panel (ESP); External Tank Umbilical (ETU); Ku-Band Deploy (KBD); Payload Bay Doors (PBD); Payload Bay Radiators (PBR); Personnel Hatches (PH); Vent Door Mechanism (VDM); and Startracker Door Mechanism (SDM). The IOA analysis process utilized available MAS hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.

Laser Shearographic Inspection for Debonds in Sprayed On Foam Insulation (SOFI)

NASA Technical Reports Server (NTRS)

Adams, F. W.; Hooker, J.; Simmons, S.

1997-01-01

Preliminary results of shearographic inspections of the test panels simulating the Space Shuttle's external tank (ET) spray on foam insulation (SOFI) are presented. Debonding of SOFI may introduce flight debris that may damage the orbiter's thermal protection system (TPS) exposing the orbiter (as well as the ET) to thermal loading. It is estimated that 90 percent of the TPS damage on the orbiter's 'belly' results from debonded SOFI during ascent. A series of test panels were fabricated, with programmed debonds of different geometries and sizes, to determine the sensitivity of shearography as a function of debond size, SOFI thickness,'and vacuum excitation. Results show that a Probability of Detection (POD) of 0.95 or better can be expected for debonds with a diameter equal to the SOFI thickness as less than 0.4-psi pressure reduction. More testing will be required to validate the laser shearography imaging process for certifying its use in nondestructive evaluation (NDE) of Space Shuttle space flight components.

KSC-2009-2010

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., an overhead crane lowers NASA's Lunar Reconnaissance Orbiter, or LRO, onto the Aronson table. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

KSC-2009-2006

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., technicians prepare an Aronson table to receive NASA's Lunar Reconnaissance Orbiter, or LRO, at left. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

KSC-2009-2005

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., technicians prepare an Aronson table to receive NASA's Lunar Reconnaissance Orbiter, or LRO, at left. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

KSC-2009-2009

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., an overhead crane lowers NASA's Lunar Reconnaissance Orbiter, or LRO, toward the Aronson table. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

KSC-2009-2004

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., technicians moved the stand with NASA's Lunar Reconnaissance Orbiter, or LRO. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

KSC-2009-2003

NASA Image and Video Library

2009-03-07

CAPE CANAVERAL, Fla. – At Astrotech in Titusville, Fla., a technician attaches cables to NASA's Lunar Reconnaissance Orbiter, or LRO. The orbiter will be rotated on the table to provide proper access for processing. The orbiter will carry seven instruments to provide scientists with detailed maps of the lunar surface and enhance our understanding of the moon's topography, lighting conditions, mineralogical composition and natural resources. Information gleaned from LRO will be used to select safe landing sites, determine locations for future lunar outposts and help mitigate radiation dangers to astronauts. The polar regions of the moon are the main focus of the mission because continuous access to sunlight may be possible and water ice may exist in permanently shadowed areas of the poles. Accompanying LRO on its journey to the moon will be the Lunar CRater Observation and Sensing Satellite, or LCROSS, a mission that will impact the lunar surface in its search for water ice. Launch of LRO is targeted for May 20. Photo credit: NASA/Jack Pfaller

Molecular and electronic structure of terminal and alkali metal-capped uranium(V) nitride complexes

PubMed Central

King, David M.; Cleaves, Peter A.; Wooles, Ashley J.; Gardner, Benedict M.; Chilton, Nicholas F.; Tuna, Floriana; Lewis, William; McInnes, Eric J. L.; Liddle, Stephen T.

2016-01-01

Determining the electronic structure of actinide complexes is intrinsically challenging because inter-electronic repulsion, crystal field, and spin–orbit coupling effects can be of similar magnitude. Moreover, such efforts have been hampered by the lack of structurally analogous families of complexes to study. Here we report an improved method to U≡N triple bonds, and assemble a family of uranium(V) nitrides. Along with an isoelectronic oxo, we quantify the electronic structure of this 5f1 family by magnetometry, optical and electron paramagnetic resonance (EPR) spectroscopies and modelling. Thus, we define the relative importance of the spin–orbit and crystal field interactions, and explain the experimentally observed different ground states. We find optical absorption linewidths give a potential tool to identify spin–orbit coupled states, and show measurement of UV···UV super-exchange coupling in dimers by EPR. We show that observed slow magnetic relaxation occurs via two-phonon processes, with no obvious correlation to the crystal field. PMID:27996007

Lunar-edge based on-orbit modulation transfer function (MTF) measurement

NASA Astrophysics Data System (ADS)

Cheng, Ying; Yi, Hongwei; Liu, Xinlong

2017-10-01

Modulation transfer function (MTF) is an important parameter for image quality evaluation of on-orbit optical image systems. Various methods have been proposed to determine the MTF of an imaging system which are based on images containing point, pulse and edge features. In this paper, the edge of the moon can be used as a high contrast target to measure on-orbit MTF of image systems based on knife-edge methods. The proposed method is an extension of the ISO 12233 Slanted-edge Spatial Frequency Response test, except that the shape of the edge is a circular arc instead of a straight line. In order to get more accurate edge locations and then obtain a more authentic edge spread function (ESF), we choose circular fitting method based on least square to fit lunar edge in sub-pixel edge detection process. At last, simulation results show that the MTF value at Nyquist frequency calculated using our lunar edge method is reliable and accurate with error less than 2% comparing with theoretical MTF value.

Spectrum/Orbit-Utilization Program

NASA Technical Reports Server (NTRS)

Miller, Edward F.; Sawitz, Paul; Zusman, Fred

1988-01-01

Interferences among geostationary satellites determine allocations. Spectrum/Orbit Utilization Program (SOUP) is analytical computer program for determining mutual interferences among geostationary-satellite communication systems operating in given scenario. Major computed outputs are carrier-to-interference ratios at receivers at specified stations on Earth. Information enables determination of acceptability of planned communication systems. Written in FORTRAN.

The Application of GIM in Precise Orbit Determination for LEO Satellites with Single-Frequency GPS Measurements

NASA Astrophysics Data System (ADS)

Peng, Dong-ju; Wu, Bin

2012-10-01

With the precise GPS ephemeris and clock error available, the iono- spheric delay is left as the dominant error source in the single-frequency GPS data. Thus, the removal of ionospheric effects is a ma jor prerequisite for an improved orbit reconstruction of LEO satellites based on the single-frequency GPS data. In this paper, the use of Global Ionospheric Maps (GIM) in kine- matic and dynamic orbit determinations for LEO satellites with single-frequency GPS pseudorange measurements is discussed first, and then, estimating the iono- spheric scale factor to remove the ionospheric effects from the C/A-code pseu- dorange measurements for both kinematic and dynamic orbit determinations is addressed. As it is known that the ionospheric delay of space-borne GPS sig- nals is strongly dependent on the orbit altitudes of LEO satellites, we select the real C/A-code pseudorange measurement data of the CHAMP, GRACE, TerraSAR-X and SAC-C satellites with altitudes between 300 km and 800 km as sample data in this paper. It is demonstrated that the approach to eliminating ionospheric effects in C/A-code pseudorange measurements by estimating the ionospheric scale factor is highly effective. Employing this approach, the accu- racy of both kinematic and dynamic orbits can be improved notably. Among those five LEO satellites, CHAMP with the lowest orbit altitude has the most remarkable improvements in orbit accuracy, which are 55.6% and 47.6% for kine- matic and dynamic orbits, respectively. SAC-C with the highest orbit altitude has the least improvements in orbit accuracy accordingly, which are 47.8% and 38.2%, respectively.

Normal orbit skeletal changes in adolescents as determined through cone-beam computed tomography.

PubMed

Lee, B; Flores-Mir, C; Lagravère, M O

2016-11-10

To determine three-dimensional spatial orbit skeletal changes in adolescents over a 19 to 24 months observation period assessed through cone-beam computed tomography (CBCT). The sample consisted of 50 adolescents aged 11 to 17. All were orthodontic patients who had two CBCTs taken with an interval of 19 to 24 months between images. The CBCTs were analyzed using the third-party software Avizo. Sixteen anatomical landmarks resulting in 24 distances were used to measure spatial structural changes of both orbits. Reliability and measurement error of all landmarks were calculated using ten CBCTs. Descriptive and t-test statistical analyses were used to determine the overall changes in the orbits. All landmarks showed excellent reliability with the largest measurement error being the Y-coordinate of the left most medial point of the temporalis grooves at 0.95 mm. The mean differences of orbital changes between time 1 and time 2 in the transverse, antero-posterior and vertical directions were 0.97, 0.36 and 0.33 mm respectively. Right to left most antero-inferior superior orbital rim distance had the greatest overall transverse change of 4.37 mm. Right most posterior point of lacrimal crest to right most postero-lateral point of the superior orbital fissure had the greatest overall antero-posterior change of 0.52 mm. Lastly, left most antero-inferior superior orbital rim to left most antero-superior inferior orbital rim had the greatest overall vertical change of 0.63 mm. The orbit skeletal changes in a period of 19-24 months in a sample of 11-17 year olds were statistically significant, but are not considered to be clinically significant. The overall average changes of orbit measurements were less than 1 mm.

In-flight performance of pulse-processing system of the ASTRO-H/Hitomi soft x-ray spectrometer

NASA Astrophysics Data System (ADS)

Ishisaki, Yoshitaka; Yamada, Shinya; Seta, Hiromi; Tashiro, Makoto S.; Takeda, Sawako; Terada, Yukikatsu; Kato, Yuka; Tsujimoto, Masahiro; Koyama, Shu; Mitsuda, Kazuhisa; Sawada, Makoto; Boyce, Kevin R.; Chiao, Meng P.; Watanabe, Tomomi; Leutenegger, Maurice A.; Eckart, Megan E.; Porter, Frederick Scott; Kilbourne, Caroline Anne

2018-01-01

We summarize results of the initial in-orbit performance of the pulse shape processor (PSP) of the soft x-ray spectrometer instrument onboard ASTRO-H (Hitomi). Event formats, kind of telemetry, and the pulse-processing parameters are described, and the parameter settings in orbit are listed. The PSP was powered-on 2 days after launch, and the event threshold was lowered in orbit. The PSP worked fine in orbit, and there was neither memory error nor SpaceWire communication error until the break-up of spacecraft. Time assignment, electrical crosstalk, and the event screening criteria are studied. It is confirmed that the event processing rate at 100% central processing unit load is ˜200 c / s / array, compliant with the requirement on the PSP.

Attitude Ground System (AGS) for the Magnetospheric Multi-Scale (MMS) Mission

NASA Technical Reports Server (NTRS)

Raymond, Juan C.; Sedlak, Joseph E.; Vint, Babak

2015-01-01

MMS Overview Recall from Conrads presentation earlier today MMS launch: March 13, 2015 on an Atlas V from Space Launch Complex 40, Cape Canaveral, Florida MMS Observatory Separation: five minute intervals spinning at 3 rpm approximately 1.5 hours after launch MMS Science Goals: study magnetospheric plasma physics and understand the processes that cause power grids, communication disruptions and Aurora formation Mission: 4 identical spacecraft in tetrahedral formation with variable size1.2 x 12 RE in Phase 1, with apogee on dayside to observe bow shock1.2 x 25 RE in Phase 2, with apogee on night side to observe magneto tail Challenges Tight attitude control box, orbit and formation maintenance requirements Maneuvers on thrusters every two weeks Delta-H Spin axis direction and spin rate maintenance Delta-V Orbit and Formation maintenance Mission phase transitions AGS support Smart targeting prediction of Spin-Axis attitude in the presence of environmental torques to stay within the science attitude Determination of the spacecraft attitude and spin rate (sensitive to knowledge of inertia tensor)Calibrations to improve attitude determination results and improve orbit maneuvers Mass properties (Center of Mass, and inertia tensor for nutation and coning) Accelerometer bias (sensitive to the accuracy of the rate estimates) Sensor alignments.