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Proximity Maneuvering of Libration Point Orbit Formations Using Adapted Finite Element Methods

Laura Garcia Taberner
Universitat Politècnica de Catalunya, April 29th, 2010
This doctoral dissertation is structured in four chapters as follows. The first chapter contains a summary of formation flying projects that have been taken into consideration since few years ago. We specially focus on the missions that have been planned to be located in a libration point regime. For completeness, this chapter also contains a general state of the art about the main reconfiguration techniques for satellite formations.
 
The main new contributions of the thesis are contained in chapters 2, 3 and 4. Chapter 2 introduces the general methodology that will be considered in all the dissertation. It is based on a discretization in time by means of a finite element approximation, and at the same time, is suitable to incorporate optimal control problems. In this chapter we study the reconfigurations using linearized equations about a nominal Halo orbit minimizing the functional given by the sum of the square of the magnitude of the maneuvers. This functional is not directly related to the fuel consumption, but has good properties concerning minimization and regularity.
 
In chapter 3 we are still working with the linearized model about the nonlinear orbit, but the functional that we optimize, given by the sum of the modulus of the maneuvers, is directly related to fuel consumption. As a consequence, the methodology can be tuned in such a way that, if possible, the user can choose to converge to bang-bang optimal controls (when possible) or to low thrust trajectories in general situations.
 
In this chapter, our objective is not only to study how the reconfigurations can be accomplished. We also consider the problem of obtaining good meshes for our finite element discretization, and up to a certain extent, to decide which is the best mesh for each kind of problem.
 
Finally, in chapter 4, we deal with non-linear and perturbed problems. In a first step we consider reconfigurations in the Restricted Three Body Problem and in a second one with JPL ephemeris. This fact slightly changes the trajectories of the spacecraft with respect to the ones obtained in the previous chapters. To correct for such deviations we design and implement a methodology based on adding small corrective maneuvers on top of the
nominal ones. We also study the magnitude of corrective maneuvers that will need to be applied in case of errors in the execution of the nominal ones. Finally, this chapter ends with some other applications that can be performed using the methodology we have developed.