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Simultaneous determinations of tesseral harmonic coefficients of gravitational field, ground station coordinates, and orbital elements by satellite photographic tracking
A data base of close approaches to the major planets has been generated via numerical integrations for a large number of planet-crossing asteroid orbits over the course of 200,000 yr; these data are then applied to such statistical theories as those of Kessler (1981) and Wetherill (1967). Attention is given to the orbits of the Toro-class asteroids, which violate the assumption of a lack of mean motion resonance locking between target planet and asteroid. A modified form of the Kessler theory is proposed which can address the problem of approaches between orbits that are either nearly coplanar or nearly tangent. A correlation analysis is used to test the assumption that the orbital elements of a planet-crossing orbit change solely due to close approaches.
An Indian Ocean satellite relay is examined. The relay satellite position is optimized by minimizing the sum of downlink and satellite to satellite link losses. Osculating orbital elements are used for fast intensive orbital computation. Integrated Van Vleck gaseous attenuation and a Crane rain model are used for downlink attenuation. Circular polarization losses on the satellite to satellite link are found dynamically. Space to ground link antenna pointing losses are included as a function of yaw ans spacecraft limits. Relay satellite positions between 90 to 100 degrees East are found attractive for further study.
New radial velocities for the classical Cepheid S Sge have been obtained and combined with previous observations to derive a new orbit. The revised orbital elements are: gamma, -10.3 +/- 0.4 km/s; K, 15.5 +/- 0.2 km/s; e, 0.23 +/- 0.02; omega, 203.1 +/- 4.2 deg; T0, 39902.3 +/- 6.6 JD; P, 675.79 +/- 0.18 days; f(m), 0.239 +/- 0.010 solar masses; a sin i, 0.935 AU = 139.9 +/- 2.0 x 10 exp 6 km; s.e., 1.2 km/s. The revised elements differ very little from the orbit determined by Herbig and Moore (1952). We have also obtained low resolution IUE spectra to search for the companion. The IUE spectra show excess flux at 1800 A when compared with spectra of the single Cepheid Delta Cep at the same (B-V)0. The spectral type of the companion determined from this flux excess is A7 V to F0 V. However, the mass of such a companion (1.7 to 1.5 solar masses) is smaller than the minimum mass (2.8 solar masses) required by the mass function and an evolutionary mass of the Cepheid. We infer that the companion is itself a short period binary.
Satellite orbital data for explorer i, vii, viii, ix and vanguard ii and iii - orbital elements
Mean atmospheric densities from changes in orbital elements of Explorer IX satellite
Orbital element fluctuations due to calculation errors and air drag, and the optimum smoothing interval for artifical earth satellites
Perturbation of Echo I satellite as explanation of orbital element changes effect on lifetime of balloon type satellites
Initial orbital elements of moon-earth trajectory for lunar dust or chunks ejected from surface
Lunar gravitational field from analysis of long term and secular variations of Keplerian orbital elements from Doppler tracking data for Lunar Orbiter 1 satellite
In a paper by the second author (Nacozy, 1981), various time elements are presented for use with the Sundman time transformation. In that paper, the time elements are given in terms of Keplerian orbital elements. We give here the corresponding time elements in terms of rectangular coordinates. Extensive references are given in the previous paper and will be omitted here. We present additional numerical experiments comparing the use of time elements and time transformations together with the use of time transformations alone. The results indicate a reduction in computational error when time elements are used.
The Navstar Global Positioning System (GPS) navigation data obtained from an experimental GPS receiver/processor package (GPSPAC) onboard the Landsat-4 spacecraft are evaluated for their accuracy and reliability for use in Landsat-4 orbit determination. Different orbit determination scenarios and methods are considered. The parameters estimated include a subset of three Landsat-4 clock parameters and an atmospheric drag coefficient, in addition to the orbital elements. It is found that Landsat-4 orbit solutions based on delta pseudorange data generally agree with the definitive solutions to the 50-meter level.
GS 0834-430, a 12.3 s accretion-powered pulsar, has been observed in seven outbursts with the BATSE large-area detectors on the Compton Gamma Ray Observatory. The first five outbursts observed by BATSE occurred at intervals of about 107 days, while the final two outbursts were separated by about 140 days. The photon energy spectrum, measured by Earth occultation in the 20 100 keV band, can be fitted by a power law with photon index alpha approximately equals -3.7 or by an exponential spectrum with temperature kT approximately equals 15 keV, with some variations within outbursts. The source has a low pulse fraction, less than or equal to 0.15 in the 20-50 keV band. We have observed significant temporal and energy-dependent variations in epoch folded pulse profiles. Because the intrinsic torque effects for this system are at least comparable to orbital effects, pulse timing analysis did not produce a unique orbital solution. However, confidence regions for the orbital elements yielded the following 1 sigma limits: orbital period P(sub orb) = 105.8 +/- 0.4 days and eccentricity 0.10 less than or approximately equals epsilon less than or approximately equals 0.17. GS 0834-430 is most likely a Be/X-ray binary.
Some simple guidelines based on the accuracy in determining a satellite formation's semi-major axis differences are useful in making preliminary assessments of the navigation accuracy needed to support such missions. These guidelines are valid for any elliptical orbit, regardless of eccentricity. Although maneuvers required for formation establishment, reconfiguration, and station-keeping require accurate prediction of the state estimate to the maneuver time, and hence are directly affected by errors in all the orbital elements, experience has shown that determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. Furthermore, any differences among the member's semi-major axis are undesirable for a satellite formation, since it will lead to differential along-track drift due to period differences. Since inevitable navigation errors prevent these differences from ever being zero, one may use the guidelines this paper presents to determine how much drift will result from a given relative navigation accuracy, or vice versa. Since the guidelines do not account for non-two-body perturbations, they may be viewed as useful preliminary design tools, rather than as the basis for mission navigation requirements, which should be based on detailed analysis of the mission configuration, including all relevant sources of uncertainty.
Some simple guidelines based on the accuracy in determining a satellite formation's semi-major axis differences are useful in making preliminary assessments of the navigation accuracy needed to support such missions. These guidelines are valid for any elliptical orbit, regardless of eccentricity. Although maneuvers required for formation establishment, reconfiguration, and station-keeping require accurate prediction of the state estimate to the maneuver we, and hence are directly affected by errors in all the orbital elements, experience has shown that determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. Furthermore, any differences among the member s semi-major axes are undesirable for a satellite formation, since it will lead to differential along-track drift due to period differences. Since inevitable navigation errors prevent these differences from ever being zero, one may use the guidelines this paper presents to determine how much drift will result from a given relative navigation accuracy, or conversely what navigation accuracy is required to limit drift to a given rate. Since the guidelines do not account for non-two-body perturbations, they may be viewed as useful preliminary design tools, rather than as the basis for mission navigation requirements, which should be based on detailed analysis of the mission configuration, including all relevant sources of uncertainty.
Some simple guidelines based on the accuracy in determining a satellite formation s semi-major axis differences are useful in making preliminary assessments of the navigation accuracy needed to support such missions. These guidelines are valid for any elliptical orbit, regardless of eccentricity. Although maneuvers required for formation establishment, reconfiguration, and station-keeping require accurate prediction of the state estimate to the maneuver time, and hence are directly affected by errors in all the orbital elements, experience has shown that determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. Furthermore, any differences among the member s semi-major axes are undesirable for a satellite formation, since it will lead to differential along-track drift due to period differences. Since inevitable navigation errors prevent these differences from ever being zero, one may use the guidelines this paper presents to determine how much drift will result from a given relative navigation accuracy, or conversely what navigation accuracy is required to limit drift to a given rate. Since the guidelines do not account for non-two-body perturbations, they may be viewed as useful preliminary design tools, rather than as the basis for mission navigation requirements, which should be based on detailed analysis of the mission configuration, including all relevant sources of uncertainty.
Monte is a software set designed for use in mission design and spacecraft navigation operations. The system can process measurement data, design optimal trajectories and maneuvers, and do orbit determination, all in one application. For the first time, a single software set can be used for mission design and navigation operations. This eliminates problems due to different models and fidelities used in legacy mission design and navigation software. The unique features of Monte 040 include a blowdown thruster model for GRAIL (Gravity Recovery and Interior Laboratory) with associated pressure models, as well as an updated, optimalsearch capability (COSMIC) that facilitated mission design for ARTEMIS. Existing legacy software lacked the capabilities necessary for these two missions. There is also a mean orbital element propagator and an osculating to mean element converter that allows long-term orbital stability analysis for the first time in compiled code. The optimized trajectory search tool COSMIC allows users to place constraints and controls on their searches without any restrictions. Constraints may be user-defined and depend on trajectory information either forward or backwards in time. In addition, a long-term orbit stability analysis tool (morbiter) existed previously as a set of scripts on top of Monte. Monte is becoming the primary tool for navigation operations, a core competency at JPL. The mission design capabilities in Monte are becoming mature enough for use in project proposals as well as post-phase A mission design. Monte has three distinct advantages over existing software. First, it is being developed in a modern paradigm: object- oriented C++ and Python. Second, the software has been developed as a toolkit, which allows users to customize their own applications and allows the development team to implement requirements quickly, efficiently, and with minimal bugs. Finally, the software is managed in accordance with the CMMI (Capability Maturity Model Integration), where it has been ap praised at maturity level 3.
Orbit stability has been thoughtfully studied in various celestial bodies. The increasing interest in Mars orbiters brings the question of the likelihood of natural decay in low altitude regimes. This paper studies the shape change of low altitude Mars orbits by carrying out large sets of numerical high fidelity simulations. Results showed that various configurations of the orbital elements gave perturbations that resulted in unstable orbits. The paper also studies the potential causes of the observed unstable regions. We computed theoretical spin-orbit resonances to study their implications in the stability at low altitudes. The resonances were tested at different initial Longitudes of the Ascending Node (LAN) and orbit inclinations to check the potential existence of latitude/longitude implications on the stability.