Stable retrograde orbits around the triple system 2001 SN263

The NEA 2001 SN263 is the target of the ASTER MISSION - First Brazilian Deep Space Mission. Araujo et al. (2012), characterized the stable regions around the components of the triple system for the planar and prograde cases. Knowing that the retrograde orbits are expected to be more stable, here we present a complementary study. We now considered particles orbiting the components of the system, in the internal and external regions, with relative inclinations between $90^{\circ}< I \leqslant180^{\circ}$, i.e., particles with retrograde orbits. Our goal is to characterize the stable regions of the system for retrograde orbits, and then detach a preferred region to place the space probe. For a space mission, the most interesting regions would be those that are unstable for the prograde cases, but stable for the retrograde cases. Such configuration provide a stable region to place the mission probe with a relative retrograde orbit, and, at the same time, guarantees a region free of debris since they are expected to have prograde orbits. We found that in fact the internal and external stable regions significantly increase when compared to the prograde case. For particles with $e=0$ and $I=180^{\circ}$, we found that nearly the whole region around Alpha and Beta remain stable. We then identified three internal regions and one external region that are very interesting to place the space probe. We present the stable regions found for the retrograde case and a discussion on those preferred regions. We also discuss the effects of resonances of the particles with Beta and Gamma, and the role of the Kozai mechanism in this scenario. These results help us understand and characterize the stability of the triple system 2001 SN263 when retrograde orbits are considered, and provide important parameters to the design of the ASTER mission.Comment: 11 pages, 8 figures. Accepted for publication in MNRAS - 2015 March 1

Optimal Trajectories for Near-Earth-Objects Using Solar Electric Propulsion (SEP) and Gravity Assisted Maneuver

The future interplanetary missions will probably use the conventional chemical rockets to leave the sphere of influence of the Earth, and solar electric propulsion (SEP) to accomplish the other maneuvers of the mission. In this work the optimization of interplanetary missions using solar electric propulsion and Gravity Assisted Maneuver to reduce the costs of the mission, is considered. The high specific impulse of electric propulsion makes a Gravity Assisted Maneuver 1 year after departure convenient. Missions for several Near Earth Asteroids will be considered. The analysis suggests criteria for the definition of initial solutions demanded for the process of optimization of trajectories. Trajectories for the asteroid 2002TC70 are analyzed. Direct trajectories, trajectories with 1 gravity assisted from the Earth and with 2 gravity assisted from the Earth and either Mars are present. An indirect optimization method will be used in the simulations

Using Upgraded Versions Of Close Approach Maneuvers As Transportation Solutions For Deep Space Missions

Gravity-Assisted maneuvers have been used as a technique to reduce fuel consumption in deep space missions for several decades now. It opened the doors of the exterior solar system. The literature shows those results, as well as new versions of this maneuver, which includes: the use of propulsion combined with the close approach, both high or low thrust; the passage by the atmosphere of a planet to help to change the trajectory of the spacecraft; the use of tethers to increase the changes in the velocity of the spacecraft. All those new versions have the goal of increasing the variations of energy given by the maneuver, making possible missions that would not be possible without this technique

Strategies For Non-Planar Configurations Of Geostationary Tethered Collecting Solar Power Satellite Systems

To collect additional solar energy during the hours of darkness and to overcome the limited Terrestrial solar power due to the diurnal day night cycle, the concept of a Geostationary Tethered Collecting Solar Power Satellite System has been proposed by several authors in the last years. This tethered system consists of a long tether used to link two bodies: a single large panel with a capability of collecting solar energy, and an Earth-pointing microwave transmitting satellite. In this manner, the solar energy would be collected directly from the space and beamed back down to any point on Earth. Planar configurations, when the panel and the microwave transmitting satellite are placed on geostationary orbits, have been usually investigated to maintain the tethered system around the Earth. However, this configuration implies that the panel and the microwave transmitting satellite must to orbit the Earth in exactly the same orbital plane of all geostationary satellites

Collision and Stable Regions around Bodies with Simple Geometric Shape

We show the expressions of the gravitational potential of homogeneous bodies with well-defined simple geometric shapes to study the phase space of trajectories around these bodies. The potentials of the rectangular and triangular plates are presented. With these expressions we study the phase space of trajectories of a point of mass around the plates, using the Poincaré surface of section technique. We determined the location and the size of the stable and collision regions in the phase space, and the identification of some resonances. This work is the first and an important step for others studies, considering 3D bodies. The study of the behavior of a point of mass orbiting around these plates (2D), near their corners, can be used as a parameter to understand the influence of the gravitational potential when the particle is close to an irregular surface, such as large craters and ridges

Low-Thrust Out-of-Plane Orbital Station-Keeping Maneuvers for Satellites

This paper considers the problem of out of plane orbital maneuvers for station keeping of satellites. The main idea is to consider that a satellite is in an orbit around the Earth and that it has its orbit is disturbed by one or more forces. Then, it is necessary to perform a small amplitude orbital correction to return the satellite to its original orbit, to keep it performing its mission. A low thrust propulsion is used to complete this task. It is important to search for solutions that minimize the fuel consumption to increase the lifetime of the satellite. To solve this problem a hybrid optimal control approach is used. The accuracy of the satisfaction of the constraints is considered, in order to try to decrease the fuel expenditure by taking advantage of this freedom. This type of problem presents numerical difficulties and it is necessary to adjust parameters, as well as details of the algorithm, to get convergence. In this versions of the algorithm that works well for planar maneuvers are usually not adequate for the out of plane orbital corrections. In order to illustrate the method, some numerical results are presented

On the use of controlled radiation pressure to send a satellite to a graveyard orbit

A very important topic in modern astrodynamics is the removal of satellites from their orbits, after the end of their missions. In this work, we propose the use of the solar radiation pressure to change the orbital energy of a satellite, to remove it from the operational region to a graveyard orbit. A mechanism for changing the area-to-mass ratio of the satellite and/or its reflectivity coefficient is used to accomplish this task. We derive an analytical model to find the maximum eccen- tricity achieved during the removal trajectory, for different initial conditions for the argument of perigee and the longitude of the ascending node. After that, the best trajectories, i.e., trajectories with low eccentricity, are integrated using a numerical model. These low eccentricity trajectories are important because satellites with disposal orbits with low eccentricity pose a lower risk of crossing the operational region during the de-orbiting.Un tema importante en la astrodinámica moderna es la remoción de satélites de sus órbitas al finalizar sus misiones. En este trabajo proponemos utilizar la presión de la radiación solar para modificar la energía orbital del satélite, y así alejarlo de la región operacional y enviarlo a una órbita “en el cementerio”. Para este propósito, se propone un mecanismo para cambiar la razón área-masa y/o la reflectividad del satélite. Obtenemos un modelo analítico para encontrar la máxima excentricidad alcanzada durante la trayectoria de remoción, para diferentes valores iniciales del argumento del perigeo y de la longitud del nodo ascendente. A continuación, las mejores trayectorias, esto es, las de menor excenticidad, se integran numéricamente. Estas trayectorias de baja excentricidad son importantes pues los satélites con ´orbitas de desecho de baja excentricidad tienen un menor riesgo de cruzar las regiones operacionales durante su eliminación.The author is thankful for the grants # 406841/2016-0 and 301338/2016-7 from the National Council for Scientific and Technological Development (CNPq); and grants # 2014/22295-5, 2011/08171-3, 2016/14665-2 and 2016/07248-6 from São Paulo Research Foundation (FAPESP).info:eu-repo/semantics/publishedVersio