Orbit Determination

Satellite-based navigation systems uses one-way ranging measurements for system orbit estimation and time-keeping, due to its operational advantage when compared with two-way ranging technique, in terms of complexity of ground monitoring stations (completely passive and requiring a simple omnidirectional antenna to track all the satellites in view). However, a sufficient number of simultaneous independent measurements is required to solve the system unknowns: in particular simultaneous visibility of multiple stations by an individual satellite (allowing to separate the ground stations clocks contributions since the SVs clocks disappear), as well as simultaneity of observations from the same monitor stations of a large number of satellites (allowing to recover the SVs clocks parameters, since the GSS clocks drop out) is the key to an effective separation in the solution of the clock contributions from the pseudo-ranges. In the Galileo IOV phase (consisting of 4 satellites on two orbital planes and a ground network of 20 Sensor Stations), the first condition is clearly fulfilled, however the second condition is not met for a considerable part of the time. If two GSSs do not see simultaneously a single Galileo satellite, they will not be able to estimate their clocks time and frequency drifts, i.e. they will not be synchronized. The free running clocks will essentially enter a holdover mode, were the relative time between the two stations will be slowly drifting as a function of the initial conditions and the stability of the clocks. The ground stations synchronization will gradually degrade with time and when a satellite will rise on the horizon they will be essentially not synchronized to the extent required to carry on a one-way-based Orbit Determination & Time Synchronization (OD&TS). During the IOV phase, the limited number of satellites available and the peculiar characteristics of the Galileo orbits will make difficult for the Orbit Determination and Time Synchronization to start producing meaningful data, therefore some form of intermediate operational configuration must be sought to help in the OD&TS process initialization. The paper will address the proposed solution to overcome the problem of Galileo system initialization starting from the intermediate configuration with first 2 satellites (first IOV Launch) up to final IOV Configuration after second IOV launch. The proposed solution will be based on a limited use of GPS to insure the synchronization of the Galileo Sensor Stations, relying on the exploitation of the Linked Common View Time Transfer (LCVTT) Technique, while the Galileo Orbit Determination and SVs clocks characterization will be carried on autonomously and independently by GPS, in a two step process, up to the achievement of the IOV Configuration with 4 satellites, when the nominal Orbit Determination and Time Synchronization process will be operated. Moreover the paper will address the development of the LCVTT Algorithm, carried on as part of development of the infrastructure aiming to support the Galileo Verification Phase currently under definition as part of Galileo Phase C/D/E1 contract. The algorithm design and implementation will be presented together with the validation carried out (both for LCVTT and MLCVTT) to verify that the synchronization accuracy is adequate to support the Galileo System Initialization.

The trajectory modeling of uncontrolled satellites close to reentry in the atmosphere is still a challenging activity. Tracking data may be sparse and not particularly accurate, the objects' complicate shape and unknown attitude evolution may render difficult the aerodynamic computations and, last but not the least, the models used to predict the air density at the altitudes of interest, as a function of solar and geomagnetic activity, are affected by significant uncertainties.

(GRAIL) mission has constructed a lunar gravity field with unprecedented uniform accuracy on the farside and nearside of the Moon. GRAIL lunar gravity field determination begins with preprocessing of the gravity science measurements by applying corrections for time tag error, general relativity, measurement noise and biases. Gravity field determination requires the generation of spacecraft ephemerides of an accuracy not attainable with the pre-GRAIL lunar gravity fields. Therefore, a bootstrapping strategy was developed, iterating between science data preprocessing and lunar gravity field estimation in order to construct sufficiently accurate orbit ephemerides.This paper describes the GRAIL measurements, their dependence on the spacecraft ephemerides and the role of orbit determination in the bootstrapping strategy. Simulation results will be presented that validate the bootstrapping strategy followed by bootstrapping results for flight data, which have led to the latest GRAIL lunar gravity fields.

Multi-sensor networks can alleviate the need for high-cost, high-accuracy, single-sensor tracking in favor of an abundance of lower-cost and lower-accuracy sensors to perform multi-sensor tracking. The use of a multi-sensor network gives rise to the need for a fusion step that combines the outputs of all sensor nodes into a single probabilistic state description. When considering Gaussian uncertainties, the well-known covariance intersection technique may be used. In the more general, non-Gaussian case, covariance intersection is not sufficient. This paper examines a fusion method based on logarithmic opinion pools and develops algorithms for multi-sensor data fusion as well as investigates weight selection schemes for the opinion pool. The proposed fusion rules are applied to the tracking of a space object using multiple ground-based optical sensors. Non-Gaussian orbit determination methods are applied to each sensor individually, and the fusion rule is applied to the combined outputs of each sensor node. It is shown that the multi-sensor fusion rule leads to an increase of nearly two orders of magnitude in the position tracking accuracy as compared to the traditional single-sensor tracking method.

The magneto-plasma sail (mini-magnetospheric plasma propulsion) produces the propulsive force due to the interaction between the artificial magnetic field around the spacecraft inflated by the plasma and the solar wind erupted from the Sun with a speed of 300-800 km/s. The principle of the magneto-plasma sail is based on the magnetic sail whose original concept requires a huge mechanical coil structure, which produces a large magnetic field to capture the energy of the solar wind. Meanwhile in the case of the magneto-plasma sail, the magnetic field will be expanded by the inertia of plasma flow to a few tens of kilometer in diameter, resulting in a thrust of a few Newton R. Winglee's group of the University of Washington originally proposed the idea of magnetic field inflation by the plasma. This paper investigates the characteristics of the magneto-plasma sail by comparing it with the other low-thrust propulsion systems (i.e., electric propulsion and solar sail), and the potential of its application to near future outer planet missions is studied.

The expanding role of positional covariance data in modern spacecraft operations leads to a need for better understanding of the time evolution of covariance by mission planners and operators. It is common to see positional covariance information presented as uncertainties in the radial, cross-track and in-track directions. While such a time history does provide some information, it tends to obscure the true directionality of the uncertainty. A methodology for improving the understanding of position covariance based on 3D visualization is presented. The methodology utilizes sequential estimation to provide insight into the evolution of the covariance in the presence of measurements and to incorporate the effects of dynamic modeling uncertainties between measurements. Display of the covariance as a triaxial ellipsoid representing a desired probability level is discussed. Various methods for interpolating positional covariance are examined resulting in a recommendation of a technique based on an eigenvalueeigenvector decomposition of the covariance.

We present a new orbit determination method, termed statistical ranging, applicable to poorly observed single-apparition asteroids having two or more observations. We examine the Bayesian a posteriori probability density of the orbital elements using a Monte Carlo technique that maps an unbiased hypervolume in element phase space. We select two observations randomly from the full set, and after introducing random deviations to the observations and generating random topocentric ranges (distances), compute a sample orbit from the Cartesian positions available. Repeating the procedure allows us to obtain a large set of orbital elements compatible with the observations. To distinguish between realistic and unrealistic orbit solutions, we can incorporate the three-or four-dimensional distribution of semimajor axes, eccentricities, inclinations, and, in some cases, the absolute magnitudes of multi-apparition asteroids as a priori information in the Bayesian technique. The application of the method to both near-Earth and main-belt asteroids clearly indicates that there may exist a broad variety of orbits, all of which are compatible with the observations. The resulting distributions of the orbital elements allow recovery attempts for lost asteroids and linkage to existing observations of single-apparition asteroids. The Bayesian approach can prove useful in constraining the broad distribution of orbits for problematic newly discovered asteroids. c 2001 Elsevier Science (USA)

The DORIS instrument on Jason-2 is the first of a new generation. The satellite receivers have now seven simultaneous measurement channels, with synchronous dual frequency phase and pseudo-range measurements. These measurements are now described in a similar manner as GPS measurements and an extension of the RINEX 3.0 format has been defined for DORIS. Data are available to users with a shorter latency.

In this paper, an efficient and simple algorithm is developed for the solution of universal Lambert's problem. The algorithm based on iterative scheme which converges for all conic motion (elliptic, parabolic, or hyperbolic). Moreover, the involved transcendental Stumpff's functions are evaluated efficiently using series and recursion formulae. Applications to the algorithm are also given.

As the spacefaring community is well aware, the increasingly rapid proliferation of man-made objects in space, whether active satellites or debris, threatens the safe and secure operation of spacecraft and requires that we change the way we conduct business in space. The introduction of appropriate protocols and procedures to regulate the use of space is predicated on the availability of quantifiable and timely information regarding the behavior of resident space objects (RSO): the basis of space domain awareness (SDA). Yet despite five decades of space operations, and a growing global dependence on the services provided by space-based platforms, the population of Earth orbiting space objects is still neither rigorously nor comprehensively quantified, and the behaviors of these objects, whether directed by human agency or governed by interaction with the space environment, are inadequately characterized. Key goals of advanced SDA are to develop a capability to predict RSO behavior, extending SDA beyond its present paradigm of catalog maintenance and forensic analysis, and to arrive at a comprehensive physical understanding of all of the inputs that affect the motion of RSOs. Solutions to these problems require multidisciplinary engagement that combines space surveillance data with other information, including space object databases and space environmental data, to help decision-making processes predict, detect, and quantify threatening and hazardous space domain activity. 1.0 INTRODUCTION This document presents an introductory overview of space surveillance, tracking, and information fusion for SDA. A relevant activity is the NATO Science and Technology Organization's Task Group (SCI-279-TG) is addressing the technical considerations for enabling a NATO-Centric Space Domain Common Operating Picture (COP). The impetus for this effort is the growing dependence by NATO and its member nations on space capabilities to achieve its mission responsibilities as well as the growing role that space, as an operational domain in its own right, is playing in matters concerning global security. NATO has recognized this important reality and increased the Alliance's collective attention on ensuring NATO operations maximize their leverage of space while ensuring the space capabilities provided by its member nations are preserved to the maximum extent possible. A critical element of ensuring the availability and efficacy of these space capabilities is the availability of a common operational perspective or picture of the space domain throughout the Alliance and its partners. The presumption is that NATO forces will be more efficient, protected and successful in their future missions if a common operational perspective can be achieved across all operational domains in which NATO must operate; air, land, sea, cyber AND space. The corollary is, without a common Alliance perspective of the space domain,

This paper develops the ontology of space objects for theoretical and computational ontology applied to the space (astronautical/astronomical) domain. It follows “An ontological architecture for Orbital Debris Data” (Rovetto, 2015) and “Preliminaries of a Space Situational Awareness Ontology” (Rovetto, Kelso, 2016). Important considerations for developing a space object ontology, or more broadly, a space domain ontology are presented. The main category term ‘Space Object’ is analyzed from a philosophical perspective, and the ontological commitments of legal definitions for artificial space objects are also discussed. Space object taxonomies are offered and space object terms are defined. Part of the taxonomy of my Space Object Ontology (SOO), a computational ontology under-development using Protege ontology editor, is introduced. Given the narrow scope of an SOO, it serves as a part or module of a broader Space Situational Awareness Ontology, or alternatively an Orbital Space Environment Ontology (Rovetto, 2016).

In this technical report, I describe the details of the code that I had written to fit a curve for a set of observational data points using the Givens Rotation method of performing QR factorization. This document is structured in a manner where I first describe what Givens Rotation is, then the complete sequence of how I used Givens Rotation to find the R factor and, finally, how I use the R factor to fit a curve that matches the measured data points.

In this paper, an efficient and simple algorithm is developed for the solution of universal Lambert's problem. The algorithm based on iterative scheme which converges for all conic motion (elliptic, parabolic, or hyperbolic). Moreover, the involved transcendental Stumpff's functions are evaluated efficiently using series and recursion formulae. Applications to the algorithm are also given.

The orbital debris problem presents an opportunity for international cooperation toward the mutually beneficial goals of orbital debris prevention, mitigation, remediation, and improved space situational awareness (SSA). Achieving these goals requires sharing orbital debris and other SSA data. Toward this, I present an ontological architecture for the orbital debris and related domains, taking steps in the creation of an orbital debris ontology. The purpose of the ontology is to capture general scientific domain knowledge; formally represent the entities within the domain; form, structure, and standardize (where needed) orbital and SSA terminology; and foster semantic interoperability and data-exchange. In doing so I hope to offer a scientifically accurate ontological representation of the orbital domain; contribute to research in astroinformatics, space ontology, and space data management; and improve spaceflight safety by providing a means to capture and communicate information associated with space debris.

The satellites of geostationary navigation overlay systems such as EGNOS (European Geostationary Navigation Overlay System) are equipped with single-frequency microwave transponders. The tracking data contain a GPS-like signal corresponding to the GPS C/A-code (Clear Access code) in the GPS L 1 -band of the electromagnetic spectrum. This signal is tracked by (some of the) commercially available GPS receivers and may be used for orbit determination and for the estimation of EGNOS clock corrections.

Arecibo delay-Doppler measurements of (99942) Apophis in 2005 and 2006 resulted in a five standard-deviation trajectory correction to the optically predicted close approach distance to Earth in 2029. The radar measurements reduced the volume of the statistical uncertainty region entering the encounter to 7.3% of the pre-radar solution, but increased the trajectory uncertainty growth rate across the encounter by 800% due to the closer predicted approach to the Earth. A small estimated Earth impact probability remained for 2036. With standard-deviation planeof-sky position uncertainties for 2007-2010 already less than 0.2 arcsec, the best near-term ground-based optical astrometry can only weakly affect the trajectory estimate. While the potential for impact in 2036 will likely be excluded in 2013 (if not 2011) using ground-based optical measurements, approximations within the Standard Dynamical Model (SDM) used to estimate and predict the trajectory from the current era are sufficient to obscure the difference between a predicted impact and a miss in 2036 by altering the dynamics leading into the 2029 encounter. Normal impact probability assessments based on the SDM become problematic without knowledge of the object's physical properties; impact could be excluded while the actual dynamics still permit it. Calibrated position uncertainty intervals are developed to compensate for this by characterizing the minimum and maximum effect of physical parameters on the trajectory. Uncertainty in accelerations related to solar radiation can cause between 82 and 4720 Earth-radii of trajectory change relative to the SDM by 2036. If an actionable hazard exists, alteration by 2-10% of Apophis' total absorption of solar radiation in 2018 could be sufficient to produce a six standard-deviation trajectory change by 2036 given physical characterization; even a 0.5% change could produce a trajectory shift of one Earth-radius by 2036 for all possible spin-poles and likely masses. Planetary ephemeris uncertainties are the next greatest source of systematic error, causing up to 23 Earth-radii of uncertainty. The SDM Earth point-mass assumption introduces an additional 2.9 Earth-radii of prediction error by 2036. Unmodeled asteroid perturbations produce as much as 2.3 Earth-radii of error. We find no future small-body encounters likely to yield an Apophis mass determination prior to 2029. However, asteroid (144898) 2004 VD17, itself having a statistical Earth impact in 2102, will probably encounter Apophis at 6.7 lunar distances in 2034, their uncertainty regions coming as close as 1.6 lunar distances near the center of both SDM probability distributions. (J.D. Giorgini).

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 Sys-tem (GPS) and the Russian GLONASS and the European Galileo systems in the future. At the Astronomical Insti-tute of the University of Bern (AIUB) LEO precise or-bit 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 numeri-cal integration techniques, as also used for dynamic orbit determination of satellites at high altitudes, are extended by pseudo-stochastic orbit modeling techniques to effi-ciently 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...

Destination: 700 km LEO - 200 km LLO
Micro Ion Thruster Optimization
Low Thrust Continuous Spiral Orbit Approach
R3BP and GEO Launch is the next step
3U/6U/12U Satellite Designs
Primary Cost Estimation
Basic Thruster Optimization:
    Initial Mass, Burn Time vs. Thrust

1] We present measurements of the density of the Martian atmosphere at $400 km altitude. Our analysis used radio tracking data to perform precise orbit determination on the Mars Odyssey spacecraft between March 2002 and November 2005. Recent improvements in a priori physical models make it possible to isolate the contribution of the atmospheric drag from the various forces acting on the spacecraft. For each spacecraft trajectory segment (arc) we adjusted an atmospheric drag coefficient (C D ), which scales the a priori model density. From the drag coefficient we obtained a time series of the measured density. These measurements at the Mars Odyssey orbiting altitude are close to noise level, and the various tests we conducted show the robustness of the measurements. We obtained a better agreement with the atmospheric model used during the second Martian year, when solar activity is lower. Using various simple exponential atmosphere models, we estimated the scale height near the spacecraft periapsis and found values between 25 and 50 km, in the lower range of expected values, and used exospheric temperature estimates to assess the role of EUV heating of the upper atmosphere. We did not observe one-to-one correlation between solar activity and exospheric density, but we detected a solar rotation periodicity in our measurements.

A force model for the lunar albedo effect on low lunar orbiters is developed on the basis of Clementine imagery and absolute albedo measurements. The model, named the De& Lzmur A&do Model 1 (DLAM-l), is a 15 x 15 spherical harmonics expansion, and is intended for improved force modeling for low satellite orbits as well as to minimise aliasing of non-gravitational force model defects in fnture lunar gravity solutions. The development of the model from the available lunar albedo data sources is described, followed by a discussion on its calibration using absolute albedo measurements and its correlation with main selenological features. DLAM-1 is next applied in low lunar orbit determination, and results for orbits typical of lunar mapping missions are presented. Finally, the effect of lunar albedo on future gravity solutions is anaIysed with particular emphasis on gravity mapping from global data sets, i.e. satellite-to-satellite tracking, which is expected to be a core experiment of coming lunar missions. It is shown that albedo-induced orbit perturbations have a magnitude and frequency signature which are non-negligible for precise orbit and gravity modeling. Radial orbit errors for low orbits are in the order of l-2 m for one week arcs. ~ODUCTION 01999 COSPAR. Published by Elsevier Science Ltd.

The first mission of NASA's New Millennium Program, Deep Space I , has, as one of its principal demonstrationtechnologies, the first autonomous optical navigation system to be used in deep space. The AutoNav system is a set of software elements that interact with the imaging, attitude control and ion-propulsion systems aboard DSI in order to accomplish optical data taking, orbit determination, trajectory correction maneuvers, ion propulsion system (IPS) control, and encounter operations. The validation of this system in the flight of DSI to Braille was very successful. Despite very substantial problems with the DSI camera, the AutoNav system was eventually able to determine the spacecraft heliocentric position to better than 200km and .2m/s, as determined by ground-based radio navigation, which for this low-thrust mission, was itself a new technology. As well as achieving this principal goal of high quality interplanetary cruise orbit determination, AutoNav successfully completed many complex and difficult operations. These operations include provision of astronomical ephemeris data to non-navigational systems onboard, planning and execution of hours-long picturetaking sessions, planning and execution of sessions of ion-engine function, planning and execution of ,trajectory correction maneuvers, and successful completion of encounter activities during the Braille approach rehearsal, and for part of the actual approach, delivering the spacecraft to as close as 2.5km of the desired impactplane-position, and starting encounter sequences to within 5 seconds of the actual encounter-relative start time. Figure 1: Diagrammatic and Comparative Description of DS1 AutoNav AUTONOMOUS OPTICAL NAVIGATION (AutoNav) for NEW MILLENNIUM DSl : Technology Validation Fact Sheet Contact: [email protected]; Jet Propulsion Laboratory, Pasadena, CA; 818-354-8724 1 CONVENTIONAL NAVIGATION DSI AUTONOMOUS NAVIGATION onboard Nnv doscd-loop target tracking. Onboard Encounter Phase: Optimum return of science wlth Rehearad tracks to within 25oLm range, knowledge error reducal to 0 5 h Imga p r o c a d 0 1 1 . Encounter Phase: GroundBPlalApp&Optlenl Nnvlgptlon. Limited In accuracy, required, .Ira reduce science. hrge flyby ranees Maneuver Computed Maneuvers: on Ground, Parameters Uplinked, requiring ground procwalng and uulysb Imaga downlinked, Nnv Comnundr developed sprcc~poritlon, VdocUy and F o r a %tlrmrlcdoaboud fi-omoptlalikt. trhngul(ba A f c u r d a b c t t u~2 5 o L m a n d a2lrmhAcbhd.

Lunar exploration introduces higher requirements in terms of performances, flexibility and cost for all spacecraft subsystems. As for the navigation, the success of GPS in providing autonomous on board orbit determination in Low Earth Orbit environment calls for a possible extension to Earth-to-Moon missions. This paper, referring to the upcoming European Student Moon Orbiter mission concept, investigates such a possibility, considering present and foreseen available GNSS signals. Carrier-to-noise levels achievable during different phases of the mission are evaluated and compared with the thresholds of current satellite navigation hardware receivers. Moreover, possible application of emerging software receivers' techniques, able to cope with weak signals, is discussed.

The application of Global Positioning System (GPS) technology to the Geostationary Earth Orbit (GEO) determination has been constrained by the poor satellite visibility and weak signal power. This situation is expected to improve when multi-constellation Global Navigation Satellite Systems (GNSS) are available in the future. This paper aims to investigate a navigation algorithm to determine the GEO state vector in real time, using multi-constellation GNSS measurements. The navigation algorithm is based on a Kalman filter where the estimated state includes position and velocity corrections to the nominal reference trajectory and clock biases. The simulation results are presented in the paper. It is concluded that this algorithm meets the requirement for autonomous GEO satellite navigation.

ABSTRACT The Altimetric Bathymetry from Surface Slopes (ABYSS), which is the proposed science payload on the International Space Station (ISS), is a Johns Hopkins University Applied Physics Laboratory-developed flight-proved delay-Doppler phase-monopulse radar altimeter capable of measuring ocean surface slope in the 6-200-km half-wavelength frequency band range with an accuracy of 0.5 mu rad , with autonomous gimbal control to compensate for the ISS structural motions. This measurement allows an improved mapping of the global bathymetry, enabling a wide range of scientific research works and applications. The nonrepeat ISS orbital ground track is ideal for ABYSS. This letter describes a simulation study on the effects of the Earth's gravity field and other errors, including thermal bending of the ISS, on the orbit determination of the altimeter instrument antenna phase center location, fulfilling the science objectives of ABYSS. Our study concluded that the error due to mean gravity field is no longer limiting due primarily to the recent Gravity Recovery and Climate Experiment gravity modeling and that the ABYSS/ISS radial orbit slope error budget in the presence of various force and measurement model errors is estimated at the 0.2-mu rad root-sum-squared (RSS) level, which satisfies the ABYSS orbit accuracy science requirement to provide an improved mapping of global bathymetry.

The beginning of the 23rd solar cycle (May 1996 to December 2008) coincided with the start of the catalogue of global ionospheric modeling using GPS data. Comparison between solar activity parameters and GPS-derived Total Electron Content (TEC) is now possible for the whole of solar cycle 23. In this study, we compared the daily sunspot number and F10.7 cm flux with the daily mean global TEC values during the entire last solar cycle. In order to better understand the ionization response, we show correlations between the daily F10.7cm delivered by NGDC-NOAA (National Geophysical Data Center - National Oceanic and Atmospheric Administration) and the daily sunspot number from SIDC (Solar Influences Data Analysis Center) with the daily mean latitudinal TEC values extracted from CODE (Center for Orbit Determination in Europe) GPS-based global ionospheric maps for the period 1995-2009. The correlations were investigated for different daily mean latitudinal ionospheric TEC: (1) expressed in geographic and geomagnetic coordinates; (2) with respect to the seasons and; (3) with respect to the different phases of the solar cycle. In general, results show in the north and south hemispheres a different ionospheric response (TEC) to solar activity (F10.7cm). Moreover, the switch from geographic to geomagnetic coordinates does not change the observed correlation between TEC and solar parameters. Finally, a larger correlation is observed at N20°-30° during the transition phase in the solar cycle.

The global navigation satellite system receiver for atmospheric sounding (GRAS) on MetOp-A is the first European GPS receiver providing dual-frequency navigation and occultation measurements from a spaceborne platform on a routine basis. The receiver is based on ESA's AGGA-2 correlator chip, which implements a high-quality tracking scheme for semi-codeless P(Y) code tracking on the L1 and L2 frequency. Data collected with the zenith antenna on MetOp-A have been used to perform an inflight characterization of the GRAS instrument with focus on the tracking and navigation performance. Besides an assessment of the receiver noise and systematic measurement errors, the study addresses the precise orbit determination accuracy achievable with the GRAS receiver. A consistency on the 5 cm level is demonstrated for reduced dynamics orbit solutions computed independently by four different agencies and software packages. With purely kinematic solutions, 10 cm accuracy is obtained. As a part of the analysis, an empirical antenna offset correction and preliminary phase center correction map are derived, which notably reduce the carrier phase residuals and improve the consistency of kinematic orbit determination results.

Optical surveys have identified a class of high area-to-mass ratio (HAMR) objects in the vicinity of the Geostationary Earth Orbit (GEO) ring [1]. The exact origin and nature of these objects are not well known, although their proximity to the GEO ring poses a hazard to active GEO satellites. Due to their high area-to-mass ratios, solar radiation pressure perturbs their orbits in ways that makes it difficult to predict their orbital trajectories over periods of time exceeding a week. To better understand these objects and their origins, observations that allow us to derive physical characteristics are required in order to improve the non-conservative force modeling for orbit determination and prediction. Information on their temperatures, areas, emissivities, and albedos may be obtained from thermal infrared, mid-wave infrared (MWIR), and visible measurements. Spectral features may help to identify the composition of the material, and thus possible origins for these objects. We have collected observational data on various HAMR objects from the AMOS observatory 3.6 m AEOS telescope. The thermal-IR spectra of these low-earth orbit objects acquired by the Broadband Array Spectrograph System (BASS) span wavelengths 3-13 mm and constitute a unique data set, providing a means of measuring, as a function of time, object fluxes. These, in turn, allow temperatures and emissivity-area products to be calculated. In some instances we have also collected simultaneous filtered visible photometric data on the observed objects. The multi-wavelength observations of the objects provide possible clues as to the nature of the observed objects. We describe briefly the nature and status of the instrumental programs used to acquire the data, our data of record, our data analysis techniques, and our current results, as well as future plans.

Space VLBI is a new observing technique which will be available in the second half of this decade by radio telescopes as interferometer elements in orbit. Besides the main astronomical interests, this new development has some very interesting potential applications in satellite dynamics, geodesy and geodynamical research which qualify space VLBI a potential new technique in space geodesy. A geodesy demonstration experiment (GEDEX) is being proposed for the Japanese space VLBI satellite-VSOP, with the main objectives of reference frame interconnections and orbit accuracy improvement of the space radio telescope. The paper gives a review of the special characteristics of space VLBI and a general background of the GEDEX proposal.

This paper discusses the motivation and general design elements of a new fundamental physics experiment that will test relativistic gravity at the accuracy better than the effects of the second order in the gravitational field strength, ∝ G 2 . The Laser Astrometric Test Of Relativity (LATOR) mission uses laser interferometry between two micro-spacecraft whose lines of sight pass close by the Sun to accurately measure deflection of light in the solar gravity. The key element of the experimental design is a redundant geometry optical truss provided by a long-baseline (100 m) multi-channel stellar optical interferometer placed on the International Space Station (ISS). The spatial interferometer is used for measuring the angles between the two spacecraft and for orbit determination purposes. In Euclidean geometry, determination of a triangle's three sides determines any angle therein; with gravity changing the optical lengths of sides passing close by the Sun and deflecting the light, the Euclidean relationships are overthrown. The geometric redundancy enables LATOR to measure the departure from Euclidean geometry caused by the solar gravity field to a very high accuracy.

Dr. Oscar L. Colombo works in the area of applications of space geodesy, including gravity field mapping, spacecraft orbit determination, and precise positioning by space techniques, mostly for the Space Geodesy Branch (Code 926) of NASA Goddard Space Flight Center. In recent years, he has developed and tested techniques for very long baseline kinematic GPS, in collaboration with groups in Australia, Denmark, Holland, and the USA. ABSTRACT

The first ESA (European Space Agency) Earth explorer core mission GOCE (Gravity field and steady-state Ocean Circulation Explorer) was launched on 17 March 2009 into a sun-synchronous dusk-dawn orbit with an exceptionally low initial altitude of about 280 km. The onboard 12channel dual-frequency GPS (Global Positioning System) receiver delivers 1 Hz data, which provides the basis for precise orbit determination (POD) for such a very low orbiting satellite. As part of the European GOCE Gravity Consortium the Astronomical Institute of the University of Bern and the Department of Earth Observation and Space Systems are responsible for the orbit determination of the GOCE satellite within the GOCE High-level Processing Facility. Both quicklook (rapid) and very precise orbit solutions are produced with typical latencies of 1 day and 2 weeks, respectively. This article summarizes the special characteristics of the GOCE GPS data, presents POD results for about 2 months of data, and shows that both latency and accuracy requirements are met. Satellite Laser Ranging validation shows that an accuracy of 4 and 7 cm is achieved for the reduced-dynamic and kinematic Rapid Science Orbit solutions, respectively. The validation of the reduced-dynamic and kinematic Precise Science Orbit solutions is at a level of about 2 cm.

First results are presented from a newly developed meteoroid orbit survey, called CAMS -Cameras for Allsky Meteor Surveillance, which combines meteor detection algorithms for low-light video observations with traditional video surveillance tools. Sixty video cameras at three stations monitor the sky above 31°elevation. Goal of CAMS is to verify meteor showers in search of their parent comets among newly discovered near-Earth objects.