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In Reference  and Chapter 3, the effect of solar and lunar gravity perturbations on a spacecraft in near-circular orbit was analyzed. However, in order to simplify the analysis, the lunar as well as the apparent solar orbits were assumed to be circular. In this chapter, this assumption is removed such that the third-body orbits are now allowed to describe Keplerian ellipses. Second and higher-order terms in the eccentricities of the lunar and apparent solar orbits are neglected in this analysis. Analytic solutions for the position and velocity components describing the perturbed motion in the Euler-Hill frame are thus obtained, generalizing thereby the results of  and Chapter 3. Once the transformations to the inertial system are carried out, all the orbital elements can be readily obtained and used for example in the maneuver planning function. This theory can be useful for the autonomous navigation of geostationary spacecraft as well as other high near-circular orbit applications such as the GPS spacecraft.
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go back to reference Kechichian, J. A. (1988). Analytic solution of perturbed motion in near-circular orbit due to luni-solar gravity in rotating and inertial Cartesian frames, AIAA Paper 88-4295-CP, AIAA/AAS astrodynamics conference, Minneapolis. Kechichian, J. A. (1988). Analytic solution of perturbed motion in near-circular orbit due to luni-solar gravity in rotating and inertial Cartesian frames, AIAA Paper 88-4295-CP, AIAA/AAS astrodynamics conference, Minneapolis.
- Effect of Luni-Solar Gravity Perturbations on a Near-Circular Orbit: Third-Body Orbit Eccentricity Considerations
Jean Albert Kéchichian
- Springer International Publishing
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