[en] The availability of full size MAPS sensors makes it possible to construct a calorimeter with pixelsize of a few tens of micrometers. This would be small enough to count individual shower particles and would allow a shower shape analysis on an unprecedented, small scale. Interesting features would be tracking capability for particle flow algorithms and a superior discrimination of single photons from neutral and charged pions at high momenta. A small Molière radius together with high transverse resolution would allow to separate close showers, induced by photons from neutral pion decay. A full scale (4 R_M, 28 X_0) prototype was constructed to demonstrate this. It features 30 micron pixelsize and a longitudinal sampling at 1 radiation length. We will show results from beam tests of this prototype at electron energies of 2 to 200 GeV

No abstract available

No abstract available

[en] 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. (paper)

[en] This paper analyzes the influence of the orbital disturbance forces in the trajectory of lunar satellites. The following gravitational and non-gravitational orbital disturbances are considered: the non-homogeneity of the lunar gravitational field; the gravitational attraction due to the third body, considering the Earth and the Sun; the lunar albedo; the solar radiation pressure. Numerical models were developed and implemented in an orbital trajectory simulator aiming to understand the dynamics of the orbital motion of an artificial satellite in lunar orbit when considering the simultaneous effect of all disturbances. Different orbits were simulated in order to characterize the major and the minor influence of each disturbing force as function of the inclination and the right ascension of the ascending node. This study can be very useful in the space mission analysis and in the selection of orbits less affected by environmental disturbances. (paper)

[en] A rendezvous mission can be divided into the following phases: launch, phasing, far range rendezvous, close range rendezvous and mating (docking or berthing). This paper aims to present a close range rendezvous with closed loop controlled straight line trajectory. The approaching is executed on V-bar axis. A PID controller and continuous thrust are used to eliminate the residual errors in the trajectory. A comparative study about the linear and nonlinear dynamics is performed and the results showed that the linear equations become inaccurate insofar as the chaser moves away from the target

[en] The objective of this work is to analyze orbital maneuvers of a spacecraft orbiting Mars, considering disturbance effects due to the gravitational attraction of the Sun, Phobos and Deimos, beyond the disturbances due to the gravitational potential of Mars. To simulate the trajectory, constructive aspects of the propulsion system were considered. Initially ideal thrusters, capable of applying infinite magnitude of the thrust, were used. Thus, impulsive optimal maneuvers were obtained by scanning the solutions of the Lambert's problem in order to select the maneuver of minimum fuel consumption. Due to the impossibility of applying an impulse, the orbital maneuver must be distributed in a propulsive arc around the position of the impulse given by the solution of the Lambert's problem. However the effect of the propulsive arc is not exactly equivalent to the application of an impulse due to the errors in magnitude and direction of applied thrust. Therefore, the influence of the thrusters’ capacity in the trajectory was evaluated for a more realistic model instead of the ideal case represented by the impulsive approach. Beyond the evaluation of the deviation in the orbital path, was considered an automatic correction of the semi-major axis using continuous low thrust controlled in closed loop to minimize the error in the trajectory after the application of the main thrust. (paper)

[en] This work aims to study and simulate the control of a spacecraft trajectory in order to correct automatically and simultaneously the orbital elements that define the orbit: semi-major axis, eccentricity, periapse argument, inclination and right ascension of the ascending node. Thus, to perform the control of the trajectory was used a propulsion system able to apply thrust with adjustable magnitude and direction of application. In this study it was considered that the propulsion system is controlled in closed loop, so the adjustments of the magnitude and direction of thrust depends on the error generated by comparing a reference state (position and velocity) and a current state. The reference state is determined according to the final orbital parameters. The current state is estimated at each step of the simulation, therefore, the reference and current states must be determined and compared at each step in order to generate the error signal that is inserted into the trajectory control system. However, the control of the orbital parameters simultaneously can be characterized as a multi-objective problem with conflicting goals. The correction of the semi-major axis causes an eccentricity modification and vice-versa. One possibility to deal with this problem is to define when and where to make adjustments for each of the parameters. Thus, the automatic control seeks the best way to correct each parameter, adjusting each one sequentially. At the end of the process all orbital parameters are automatically adjusted and maintained due to the use of the closed loop control system

[en] The asteroid 243 Ida located in the asteroid belt, between Mars and Jupiter, is the fourth largest asteroid of the Koronis asteroid family, with an average diameter of 31.3 km and a mass around 4.2×10¹⁶ kg, and a small moon, Dactyl. In order to study the dynamics of this system, orbital trajectories are simulated around Ida considering, besides the gravitational attraction of Dactyl, the non-central gravitational field of the asteroid, defined by a polyhedral model that defines the shape and the non-uniform mass distribution of the body. In this way, the magnitude and the behaviour of such forces, and also their influence on the orbital elements that define the trajectory of the space vehicle, are evaluated and analysed. (paper)

[en] 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. (paper)