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The Brazilian Colloquium on Orbital Dynamics (CBDO) is a scientific meeting that takes place every two years since 1982. Its main goal is to bring together researchers, professors, undergraduate and post graduate students who develop works on pure and applied Celestial Mechanics in the same lectures and discussions atmosphere, and thus to drive collaboration between these areas. The XIX CBDO 2018 took place at the Test and Integration Laboratory (LIT) of the National Institute for Space Research (INPE), in São José dos Campos (SP), Brazil, between December 03 - 07, 2018, and more than 230 attendant from South and North America, Europe and Asia presented works of which 31 were selected for publication in this volume of the Journal of Physics: Conference Series. The highlight of this issue is the consolidation of the research groups on Celestial Mechanics in several Brazilian universities.

All papers published in this volume of Journal of Physics: Conference Series have been peer reviewed through processes administered by the proceedings Editors. Reviews were conducted by expert referees to the professional and scientific standards expected of a proceedings journal published by IOP Publishing.

In March 2015 the European Space Agency's Rosetta spacecraft performed a close flyby over the surface of comet 67P/Churyumov-Gerasimenko of just 15 km from the comet center. This comet belongs to the Jupiter family with an aphelion at 5.5 au and a perihelion at 1.25 au. Its orbital period is 6.5 years. The ROSINA (Rosetta Orbiter Sensor for Ion and Neutral Analysis)/DFMS (Double Focusing Mass Spectrometer) instrument on board Rosetta reported the first detection among others of glycine (C₂H₅NO₂) an amino acid, other prebiotic molecules, phosphorus atom (P), and fluoromethylidyne (CF), which is not a stable chemical species but a metastable radical [1]. In the case of phosphorus, the search for the parent (PH₃, PH, PO, PN, CP, HCP, HPO, CCP) was unsuccessful although these species have been detected mostly in the interstellar medium [2]. This work reports a study of model-dependent chemical networks, based on several databases such as UMIST and NIST among others, to explain the formation of CF radical in comets from gaseous tetrafluoromethane (CF₄).

The Geostationary Operational Environmental Satellite - GOES 16 has an arrangement of planar antennas for uplink and downlink communications in the L-band range, UHF frequency (1694,3/1694,9 MHz), 16 W transmission power and a planar antenna array gain of 28 dBi, for communication operations and data scanning. The structure of antennas in the form of arrangement combines high directivity in the electromagnetic signal and reduction of the broadside, which corresponds to a smaller angular variation. Assuming a communication channel for the GOES 16, from the phase arrangement of planar antennas using 4 rectangular radiant elements of 30 cm × 30 cm (patch) in the transmission (downlink), we defined an expression for the gain of the array of planar antennas and model an acceleration for the satellite, due to the effect of the electromagnetic perturbation it admits, antenna theory and the energy-momentum conservation laws. For a state vector - 04/02/2019, at 18h 40m 12.44s, we implemented a routine using numerical methods with the equation of movement in the form of Cartesian components, which can be used for both keplerian movement as well as adding the desired disturbing accelerations. We propagate its orbit over a period of 5 days, with a step of 10 minutes, and correlate the results of this propagation in the propagated orbital model without disturbance and with the disturbance of the acceleration on the satellite of electromagnetic origin, centered in the phase arrangement of flat antennas. The perturbative effect of this model is applied on GOES 16 taking into account satellite mass, antenna characteristics, radiated power and maximum antenna gain. The numerical integrator used for the solution of the satellite motion equation is based on the fourth and fifth degree Runge-Kutta method, and the results shows that the phase arrangement of planar antennas with the described configuration implies a significant electromagnetic disturbance, changing the components in the direction (radial, transverse and normal) and the coordinates XYZ.

This work investigates an alternative strategy to exploit future communications satellite generations including a final stage of lunar observations. For that, we explore impulsive transfers between geostationary orbits and lunar gravitational capture orbits in a full 4-body dynamical model with the Sun, Earth, Moon and spacecraft. Criteria to seek natural transfer orbits between the geostationary orbit and the vicinity of the Moon are defined considering escape properties of trajectories of the Circular Restricted Three Body Problem (CR3BP) as a guide. Namely, we select initial conditions of the 4-body model with energies that favors Earth-Moon transfers that remain around the Moon for a long time. As a case of study, we selected the current Brazilian geostationary satellite Star One C4. After a broad analysis of initial conditions and their transport behavior, we select potential transfers that reaches a near vicinity of the GEO orbit with an sufficiently small inclination with respect to the terrestrial equator. Time evolution of candidate solutions are analyzed and Δν budget and propellant mass are computed. As well as some current proposals for space debris mitigation, our strategy requires that additional mass of propellant besides onboard propulsion systems to perform final maneuvers have to be foreseen in the design of future generations of these communication satellites.

The dwarf planet Haumea is a very interesting celestial body due to the characteristics of its physical form and also the recently observed ring. A Kuiper Belt object, Haumea is a triaxial ellipsoid with dimensions of approximately 513 × 852 × 1161 (km), with a mass of 4.006 × 10²¹ kg and a rotation period of 3.915341h. The dwarf planet Haumea has its system formed by two natural satellites, the moons Namaka and Hi'iaka. We have presented an analysis of orbits around the dwarf planet Haumea taking into account the influences of the perturbations of its nonsphericity (J₂, J₄, C₂₂). We have found that the C₂₂ term and the rotation rate of Haumea have contributed strongly to reduce the variation rate of the periapsis radius of the spacecraft. We have calculated the spherical harmonics of Haumea taking into account the most current values for the semi-axes of the ellipsoid and we have presented a comparison with the values of the harmonics found in other works.

In this study, the Weighted Essential Non-Oscillatory (WENO) scheme is implemented for numerical approximations of the pressureless gas dynamics equations. The Harten-Lax-van Leer-Contact (HLLC) approximate Riemann solver serves as a basic building block with the fifth order WENO scheme involving the first order Lax-Friedrichs scheme as a flux limiter for the numerical stability. In particular, the WENO scheme suppressed the oscillatory around the delta shock waves and the flux limiter guaranteed the positivity of densities around the vacuum. Lastly, numerical testes for one-dimensional problems are presented for capturing the delta shock waves and vacuum.

This paper aims at solving motion equations for a Spatial Tether System, composed by a principal satellite and a sub-satellite, through an initial know condition. The translational motion of the sub-satellite around the principal satellite is described in spherical coordinates, described by distance and the angles that position the vector between both satellites. The rotational motion of a sub-satellite S₂ is described by Euler'equations and the cinematic equations for 3-2-3 Euler angles. The results of dynamic propagation show that the sub-satellite moves around the principal satellite in a precession anticlockwise motion, and that also vertically oscillates throughout this motion, with an amplitude of approximately 10°, for the adopted conditions. The numeric propagator of sub-satellite trajectory throughout the principal satellite and rotational motion of sub-satellite can be adapted for other types of Space Tether Systems.

This work investigates the stability of the equilibrium points that occur around the asteroid (21) Lutetia, assuming that this body has a constant velocity of rotation and is immersed in a gravitational field, whose force of attraction presents a perturbation with respect to the central force due to the irregular mass distribution of the asteroid. For the calculation of the potential, as well as of the effective potential, was used the method of the expansion of the potential in series, associated to the asteroid decomposition in tetrahedral elements. The zero velocity curves for a massless particle orbiting the gravitational environment were analyzed. The linearized dynamic equation in the vicinity of the equilibrium points, the associated characteristic equation, and the Jacobi constant were calculated. The validation of the results was ratified by simulations of trajectories around these equilibrium points, considering the gravitational field modelled. It should be emphasized the general nature of the procedures adopted in this work, that is, they can be applied to any other asteroid.

The energy of a spacecraft, relative to the primary body of the system, before and after a Powered Swing-By maneuver with an impulse applied during the close encounter with the secondary body is studied. The Powered Swing-By maneuver is a combination of the effect of the gravity of a celestial body and an impulse applied to the spacecraft during its passage by the periapsis of its orbit relative to the secondary body. This combination modifies the spacecraft's trajectory, changing its parameters and, consequently, its energy. The objective is to quantify the effect of different mass parameters on the optimum direction to apply the impulse and in the energy variation of this more complex maneuver. It is focused on the two-dimensional and elliptical maneuver. Optimum solutions for energy gains are presented.

In this paper we describe a numerical model of a helicon plasma thruster device currently under development at the Laboratory of Plasmas at the University of Brasilia. The two-dimensional model is obtained by representing the device using a cylindrical geometry, and neglecting variations in the azimuthal direction. The computational code solves the electrostatic equations and the equations of motion of charged particles self-consistently using the particle-in-cell approach. Our model predicts the emergence of a structure similar to a current-free double-layer in the plasma, which has been detected in similar experiments.

The use of innovative technologies in space missions has evolved considerably in the last decade. The use of large cables in space structures to connect spacecrafts and satellites with the goal to minimize cost of missions created a new field to be explored. A brief explanation will be considered about papers related to the equilibrium and stability of the movement of space systems connected by cables, known as Tether Systems. It will be presented the mathematical formulation for the system formed by two point masses connected by a tether in the central force field, in a Keplerian movement. The Lagrangian formulation was used to describe the rotational movement of the dumbbell-like system. Results of system behavior, tension and kinetic energy will be presented for two different situations, considering equal masses and different masses.

It is estimated that more than 22.000 objects are in orbit around the Earth, with a total mass of 8.400.000 kg. These numbers consider only objects with dimensions above 10 cm and some non-operational, but still orbiting satellites without control (debris). The debris represent a hazard to operational satellites and aerospace operations due to the high probability of collisions. Due to the interaction of the debris with the atmosphere of the Earth and the solar activity, they began to lose energy and finally decay. During the de-orbit process, the debris fall into the Earth's atmosphere at hypersonic speeds and these objects can break-up and/or be fragmented due to the aerodynamics, thermal and structural loads. It is important to obtain the trajectory and attitude of any fragment to determine the possible survival mass, impact area, hazard conditions and risks to the population, the air traffic control, and infrastructure. In this case, it is implemented a computational code to integrate the equations of motion to propagate the dynamics and kinematics of spherical debris or propellant tanks. It is also analyzed the results of trajectories with six degrees of freedom, atmospheric winds, and Magnus effect. A voxel method is implemented to analyze the tanks heat transfer, surface temperature and structures stress. To determine and observe the influence of the rotation and the Magnus force in six reentry spherical bodies, three materials are selected; aluminum alloy, due to its application in many aerospace structures; titanium and graphite epoxy I, due to their highest melting point and specific heat. Generally, these materials are used in tanks and rocket motors. More than 62 trajectories were simulated. The mathematical model and computational code were validated in three degrees of freedom. Results are compared with data from other computational tools available in the scientific literature. The results show a good approximation with reported cases of study. New results are generated in the simulations of rotational bodies, due to the influence of aerodynamic forces in the trajectory and the changes in the stagnation regions. Due to the implementation of wind and rotation of the debris, the fragments increased the survivability and the dispersion area.

This paper presents the study of the influence of the oblateness of one planet in the temporary gravitational capture of a particle (or a spacecraft) around the main celestial body of the system. The gravitational capture is a physical phenomenon where a celestial body (or a spacecraft), initially in a hyperbolic orbit with positive energy with respect to the main body of the system, is captured by another celestial body, changing the energy of the orbit to negative during a determined time. In this way, we have a temporary capture of this body or spacecraft. In the case of a spacecraft, it is possible to make this capture permanent by the application of a propulsion system to the spacecraft. This technique is done with savings over the traditional maneuvers performed without the support of the temporary capture, so it can help mission designers in the planning of the mission. As an example, the system used in the present paper is composed by the Sun-Jupiter-particle, and this system is modeled by the circular restricted three-body problem (CR3BP) plus the oblateness of Jupiter. To determine the effects of the oblateness in the dynamics of the particle, it is used the design of experiments to generate the data for a global analysis of the model. The oblateness of the planet is varied to show the effects of this parameter.

In this work we compare the performance of two types of Hall thrusters, namely, the SPT-100 and the PHall-IIc, the latter being developed by the Plasma Physics Laboratory at the University of Brasilia. The comparison is carried out by performing particle-in-cell simulations. We compute thruster parameters such as thrust and specific impulse. Our results can contribute to the design of electric propulsion devices for future Brazilian space missions.

Jupiter is a tempting body as a target for a space mission. Since this planet owns a complex system of moons, many discoveries and analyses can be done. Its Galilean moons, Io, Europa, Ganymede, and Callisto, are also of interest, especially Europa with the intriguing hypothesis of the oceans below its icy surface. This work aims to study the perturbations on a spacecraft in an orbit around Europa, the gravitational attraction of the Sun, Jupiter, Io, Ganymede, Callisto, and the gravitational potential of Europa are considered.

The present work presents a test of two Hamiltonians that produce integrable models recently proposed to study the roto-orbital motion of an axisymmetric rigid body in motion under a central gravitational field. The dynamics assumed here is approached by the motion of an axisymmetric rigid body orbiting another massive spherical one. Based on the concept of intermediary, both models are treated in Hamiltonian formalism, as perturbation of the Keplerian-Eulerian motion, using canonical variables associated to the total angular momentum. An analysis of parameters introduced to visualize possible different applications are made, in this case with special focus in binary asteroid type dynamics. The parameters space analysis present comparisons of two recently proposed intermediaries with respect to the original non-analytically integrable model and with respect to each other. In conclusion, both models behave well in regions of the parameters space where they were proposed to be valid.

This work aims to study the main perturbative effects on SGDC - Geostationary Satellite of Defense and Communication, result of the partnership between Telebrás and Brazilian Ministry of Defense. The name "SGDC" corresponds to the group of satellites that will be launched with the purpose of transmitting broadband Internet to the less favored zones of Brazil and to intermediate communication between the military sectors. Through numerical simulations, analyses are carried out to determine the influence of the perturbations of the terrestrial gravitational field, the solar radiation pressure and the lunissolar gravitational force, besides checking the combination of these perturbations, evaluating which of the effects presents predominance.

The main goal of the present paper is to show how to form a constellation of small satellites when all small satellites leave from a larger spacecraft. It is assumed that they will all stay in a planar formation, in the same orbital plane of the mother spacecraft, just dispersed in terms of mean anomaly. Initially, it is considered that the larger satellite orbits the Earth in an almost circular orbit with 2000 km of altitude. Then, a search for initial conditions of the small satellite is performed to find the best ones to move it away from the larger satellite such that it is allocated in a co-orbital orbit with respect to the larger satellite. These initial conditions should minimize the consumption of an impulsive maneuver required to move away the small satellite from the large one to put it in course to its final orbit, trying to make the best use of the most relevant perturbations, such as the solar radiation pressure and the oblateness of the Earth. In a second study, we will analyze the influence of the solar radiation pressure, depending on the A/m ratio of the spacecraft, in the trajectories.

Saturn currently has about 62 moons already discovered. This number is uncertain due to the numerous of objects that orbit the planet, but it guarantees the placement of the second planet with the highest number of moons in the Solar System, each one with diverse physical and orbital characteristics. Among all the large moons of Saturn, Dione and Enceladus are the subject of the present work, whose objective is to evaluate strategic trajectories aiming to approach an artificial satellite to these moons, in a near-regular cadence. In the vicinity of Dione and Enceladus, the artificial satellite is significantly perturbed by the gravitational potential of Saturn, which in this work is expanded in spherical harmonics (J2), and also by the gravitational field of the 13 largest moons. All simulations were done using the Spacecraft Trajectory Simulator, an orbital simulator capable to consider continuous propulsion, and trajectory control in closed loop.

Near Earth objects (NEOs) are comets and asteroids that orbit the vicinity of the Earth. The scientific interest in comets and asteroids is due in large part to their status as remnants remaining relatively unchanged from the process of forming the solar system some 4.6 billion years ago, and some asteroids are made of precious metals such as platinum. Genetic algorithms are a particular class of evolutionary algorithms that use techniques inspired by evolutionary biology such as heredity, mutation, natural selection, and recombination. They can also be defined as global optimization algorithms and model a solution to a specific problem. This method provides a way to find solutions to problems that would be unlikely to be analytically feasible. The present work aims to use this method to optimize the consumption of fuel in Rendezvous maneuvers in interplanetary missions in a context where one wishes to send a probe to an asteroid to study it.

The objective of this work is to validate the GSAM propagator using new data provided by the National Institute for Space Research (INPE) from SCD1 and SCD2 data collection satellites, with emphasis on long interval simulations without daily data updates. Originally, only 40 days of data were available to test the program, constraining any attempts to measure its precision more accurately. Recently, over two decades of data regarding both satellites' orbital and attitude parameters were provided, allowing further studies and validation of the program. The rotational motion equations are composed by the gravity gradient torque, aerodynamic torque, solar radiation pressure torque, residual and eddy current magnetic torques, the latter using a dipole geomagnetic model. The results are considered fitting when the mean deviation between the calculated variables and the real satellite data stay within 0.5° for the right ascension and declination angles and 0.5 rpm for the spin velocity. Intervals that meet the required precision were found for all years, from three to up to 15 days of simulation without data update. The consistent detection of such intervals further corroborate the use of the propagator to estimate the orientation of the satellites studied in their missions.

For many spacecraft to accomplish their mission goals there is a need to have an accurate control of their spatial orientation (attitude), so that instruments aboard the spacecraft, such as remote sensing cameras, telescopes, and other scientific payload be correctly pointed to their targets. For that purpose, many spacecraft employ sophisticated attitude control systems, fed by a variety of attitude sensors: sun sensors, horizon sensors, gyroscopes, star trackers, etc. Among the attitude sensors capable of providing an absolute attitude measurement, star trackers are considered the most accurate. This work presents an overview of the test and calibration infrastructure built for an autonomous star tracker (AST) being developed by the Electro-Optics group of INPE, result of an effort to increase the competence of the country in attitude control systems.

In face of the growing exploration of space by humanity, at first this work aims to synthesize a little information from Space Law based on legislations and discussions on the subject. Going further, the main focus of the project are the space debris and its mitigation guidelines of the United Nations Committee on the Peaceful Uses of Outer Space. The possible consequences of non-compliance with these standards for orbits around the Earth have been investigated and the scenario at the date of approval of the document implementing the guidelines at the UN has been compared with the current scenario in order to verify the impact of the measures summarized therein on the missions launched since then.

Dynamical studies suggest that most of the circumbinary discs (CBDs) should be coplanar. However, under certain initial conditions, the CBD can evolve toward polar orientation. Here we extend the parametric study of polar configurations around detached close-in binaries through N-body simulations. For polar configurations around binaries with mass ratios q below 0.7, the nominal location of the mean motion resonance (MMR) 1 : 4 predicts the limit of stability for e_B > 0.1. Alternatively, for e_B < 0.1 or q ∼ 1, the nominal location of the MMR 1 : 3 is the closest stable region. The presence of a giant planet increases the region of forbidden polar configurations around low mass ratio binaries with eccentricities e_B∼ 0.4 with respect to rocky earth-like planets. For equal mass stars, the eccentricity excitation Δβ of polar orbits smoothly increases with decreasing distance to the binary. For q < 1, Δβ can reach values as high as 0.4. Finally, we studied polar configurations around HD 98800BaBb and show that the region of stability is strongly affected by the relative positions of the nodes. The most stable configurations in the system correspond to polar particles, which are not expected to survive on longer time-scales due to the presence of the external perturber HD 98800AaAb.

Observation of stellar occultation by objects of the Solar System is a powerful technique that allows measurements of size and shape of the small bodies with accuracies in the order of the kilometre. In addition, the occultation star probes the surroundings of the object, allowing the study of putative rings/debris or atmosphere around it. Since 2009, more than 60 events by trans-Neptunian and Centaur objects have been detected, involving more than 34 different bodies. Some remarkable results were achieved, such as the discovery of rings around Chariklo and Haumea, or the high albedo of Eris, the lack of global atmosphere around Makemake and the discovery of the double shape of 2014 MU₆₉, among others. After the release of Gaia catalogues, predictions became more accurate, leading to an increasing number of successful observations of occultation events. To keep track of the results achieved with this technique, we created a database to gather all the detected events worldwide. The database is presented as an electronic table (http://occultations.ct.utfpr.edu.br/), where the main information obtained from any occultation by small outer solar system objects are listed. The structure and term definitions used in the database are presented here, as well as some simple statistics that can be done with the available results.

We are going to present preliminary experimental results for our latest PHALL thruster of the annular type, which has been greatly improved with a new magnetic circuit design that allows operation with the magnetic field perpendicular (normal configuration) or parallel (magnetic shielding) to the thruster walls; where the benefit of this last magnetic field configuration is an improvement of three orders of magnitude on the thruster's lifetime. For the first time, we will report on PHALL II-C operation up to 620 W with the generation of up to 41.39 mN of force and 2286.22 s of specific impulse, using a hollow cathode. These tests elevate the TRL of our thruster to TRL 4, with more realistic tests planned in our near future. We will also detail current work on the control and dissipation of thermal energy which will allow our thrusters to operate for long periods of time in space.

Hall thrusters are one of the most successful electric thrusters for space application that has been developed until now. The Plasma Physics Laboratory of the University of Brasília (UnB) has been developing a Permanent Magnet Hall Thruster (PHALL) for the Brazilian Space Program since 2004. Recently we have achieved important experimental results satisfying our initial goals of generating a force above 40 mN with powers around 620 W. We will discuss in this article possible applications of this thruster to nano and microsatellites with powers above 50 W. Meanwhile, a complete description is given of our present and future installations where the new thruster will be tested; taking advantage of our new 1.5 m diameter vacuum chamber (the old chamber had 0.5 m in diameter), which intends to test our thruster in the most realistic conditions, including mounting and testing on a 3U CubeSat structure, which is where we intend to start testing our thruster in a real mission in space.

Artificial satellites in low Earth orbit have as main disturbance the atmospheric drag, which is a non-conservative disturbance that causes the satellite to lose orbital energy due to the friction with the air. Basically, the drag force is a function of the velocity, the local air density and the satellite's constructive parameters. The air density is a function of altitude, longitude, latitude, geomagnetic index and solar activity. Solar storms are responsible for a wide range of terrestrial effects, especially in damage to telecommunications systems. Another relevant effect of solar activity is the variation in the volume of the atmosphere and consequently in the value of the air density for a given altitude, longitude and latitude. This work provides an initial approach, through simulation, in the engineering effort to deal with this disturbance.

This work presents a study related to the perturbation of the Sun in the orbital evolution of artificial satellites on frozen orbits around Mars. The effect generated by the non-sphericity of Mars is considered. Frozen orbits are orbits that have the eccentricity and the argument of periapsis of the nominal orbit constant on average. Our goals are to verify how the gravity effect of the Sun affect such orbits and study their stability as a function of the parameters involved. We have performed numerical integrations considering different initial conditions for the frozen orbits. Using the method of the integral of the disturbing accelerations it is possible to calculate the magnitude of the total velocity variation due to each perturbation, as well as the degree and order of the most significant spherical harmonic over a long time. Our results showed that the effect of the perturbation of the Sun on the satellites depends on the initial values of eccentricity and semi-major axis of the orbit of the satellite.

Within the framework of the Restricted Three Body Problem (RTBP), we consider the orbital evolution of a circumbinary (CB) planet. We develop a simple analytical model to explain the mean behaviour of the planetary eccentricity and better identify the nature of the main contributors of eccentricity oscilations. Our theory is validated by comparing with N-body simulation. We analyse the effect of including dissipative forces in our theory and compare with a N-body simulation in which the dissipative force is the interaction with a protoplanetary disc. Finally, we use our model to explain dynamical maps in the CB region and distinguish domains where the secular dynamics dominate over the resonant perturbations.

The study of evasive maneuvers of spacecraft in the face of the danger of collision with space debris is essential and important today due to the growth of missions at low altitudes (LEO). Bodies that orbit this region are affected by the Earth's drag force. In this work, we present the results of numerical simulations for collision dynamics between spatial objects (operational vehicle and space debris) and investigate evasive maneuvers of the vehicle under the possibility of collision with a debris in atmospheric drag environment. We tested the evasive maneuvers for collision conditions in this environment using the exhaust velocity of the vehicle's propulsion system. We find three important results: 1) evasive maneuvers are possible and the atmospheric drag can contribute to the removal of the collision objects in small and medium intervals of time; 2) for very small initial velocities, there are natural quasi-Rendezvous, under the action of gravity alone and drag; 3) propulsion in conjunction with drag can enable evasive maneuvers for various spatial debris sizes.

Specific application of satellites require the building and maintenance of orbits with high altitude. In these cases, conventional orbital transfer methods may prove to be infeasible, especially with respect to fuel consumption. In this work we present an alternative strategy for an orbital transfer of a spacecraft from a low circular orbit to a high and retrograde orbit in the Earth-Moon System. Considering a four-body scenario, the technique consists in the use of appropriate perturbations applied around unstable periodic orbits predicted in the Restricted Three-Body Problem (R3BP), associated with the exploration of the sensitive dependence of the transition regions between capture and escape in the Earth-Moon System. In addition to small transfer times, the results obtained show that it is possible to obtain stable high and inclined final orbits with a low cost in terms of increment of velocity.