Orbit Determination Error Research Articles

Precise tracking and orbit determination of ERS-1

The application of the ERS-1 altimeter for investigating the global ocean circulation requires that the satellite's orbit, and in particular the radial position component, are known very accurately. Results are presented of orbit determination error analyses for 15 min, 2 revolution and 3 day data arcs, applying laser, TRANET and PRARE tracking systems. For the center part of the short arc radial orbit errors of less than 10 cm are achievable. For the multirevolution arcs the global rms radial error is found to be about 0.6 m and is dominated by the gravity field model error contribution. Finally, the feasibility of applying the altimeter as a tracking device is discussed and orbit determination results are presented from the processing of actual SEASAT altimeter data.

Orbit determination error analyses for POPSAT and ERS-1

The missions of both the future ESA geodetic POPSAT and remote-sensing ERS-1 satellites require that the orbits are known very accurately. This paper summarizes some results of POPSAT and ERS-1 orbit determination error analyses for 1–5 day arcs of tracking data acquired by global tracking networks, and of error analyses to estimate the accuracy with which the positions of ground stations and of the earth's poles can be determined from POPSAT tracking data. In addition, the accuracy of tailored gravity models for POPSAT is addressed and some simulation results are summarized. Finally, the application of POPSAT-to-ERS tracking data is discussed and it is shown that this is a very-attractive tracking concept for a follow-on spacecraft in the ERS-family.

Voyager orbit determination at Jupiter

This paper summarizes the Voyager 1 and Voyager 2 orbit determination activity extending from encounter minus 60 days to the Jupiter encounter, and includes quantitative results and conclusions derived from mission experience. The major topics covered include an identification and quantification of the major orbit determination error sources and a review of salient orbit determination results from encounter, with emphasis on the Jupiter approach phase orbit determination. Special attention is paid to the use of combined spacecraft-based optical observations and Earth-based radiometric observations to achieve accurate orbit determination during the Jupiter encounter approach phase.

Orbit Trim Strategies for the 1975 Mars Viking Mission

Propulsive maneuver strategies have been developed for the Mars orbital phases of the Viking mission, including the activities before and after lander deployment. These strategies have been formulated as fixed sequences of orbit parameter-correction maneuvers which will be performed on specified spacecraft revolutions. Certain adaptive features are also available to ensure efficient use of propellants. Numerical simulations demonstrate that it is possible (to within a very high probability) to satisfy mission requirements on orbital navigation, using these strategies and Viking technology. HE United States will send two unmanned spacecraft to Mars during the 1975 launch opportunity. Each of these spacecraft will orbit Mars and deploy a soft-lander to the surface of the planet as part of the Viking Project. These objectives place a number of interesting and stringent requirements on the control of the satellite orbit to obtain near periapsis reconnaissance of the landing site and to prepare for lander release. As a result, the preflight navigation analyses and maneuver strategy development have been considerably increased in scope and complexity over previous missions. Reference 1 describes the Viking mission objectives and overall navigation profile from trans-Mars injection through the postlanding station-keeping phase. During the Mars orbit trim (MOT) phase of each flight, a sequence of orbit trim spacecraft propulsive maneuvers is performed to remove the effects of orbit determination and maneuver execution errors, plus any intentional biases, remaining after earlier maneuvers in the flight. The sequence of trim maneuvers is performed according to a predetermined strategy designed to achieve the mission objectives. This paper describes the maneuver objectives and the resulting strategies for the Mars orbital phases of the Viking mission, including the activities before and after lander deployment. The prelanding activities are treated in Sec. II-IV, with the consideration of postlanding activities being reserved for Sec. V. Numerical results are included in Sec. IV and V to demonstrate that the strategies satisfy all mission requirements which have been identified The preflight navigation analyses for Viking have included the analyses of the two typical missions defined in Table 1. The . typical candidate for the first spacecraft is referred to in this paper as Mission 1, while the candidate for the second spacecraft is called Mission 2. These example missions are considered in this paper to clarify techniques and to provide realistic numerical results. They are not intended to contribute a parametric analysis of all possible Viking missions. Table 1 Selected parameters for Missions 1 and 2 Parameter Mission 1 Mission 2

Maneuver Design and Implementation for the Mariner 9 Mission

The maneuver strategy and operational techniques employed in controlling the Mariner 9 flight path from earth launch, through interplanetary space, Mars orbit insertion, and the subsequent orbital trim maneuvers are presented. It is shown how the maneuver strategy was tailored to meet the mission requirements with maximum reliability in the presence of launch vehicle injection, orbit determination, and spacecraft maneuver execution errors as great as 3 sigma. The major error sources and constraints are discussed. The in-flight results are summarized and are compared with the preflight predictions.

Errors in orbit determination

The gaussian noise which affects tracking measurements causes an error in the computed value of the orbit parameters. This study provides a method for evaluating: (a) the length of the arc over which the satellite must be tracked; (b) the number of measurements to be made along this arc; (c) the position of the arc with respect to the orbit, necessary to reach the desired accurary of the calculated orbit parameters for a given pointing error of the tracking antenna. It has been assumed that the errors of the tracking measurements have a known gaussian probability distribution, which may differ for each measurement. The equations relating the orbit parameters to the measurements performed have been linearized. It has been shown that the orbit parameters are gaussian random variables and their variance has been calculated as a function of (a), (b) and (c).