Papers by evandro marconi rocco
Gravitational disturbances generated by the Sun, Phobos and Deimos in orbital maneuvers around Mars with automatic correction of the semi-major axis
Journal of physics, Oct 7, 2015
The objective of this work is to analyze orbital maneuvers of a spacecraft orbiting Mars, conside... more 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.
Simulation of optimal impulsive maneuvers of a spacecraft in orbit of Mars considering constructive aspects of the propulsion system
Journal of Physics: Conference Series, 2013
This work aims to study and simulate the control of a spacecraft trajectory in order to correct a... more 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: semimajor 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.
Journal of Physics: Conference Series, 2019
This work investigates the stability of the equilibrium points that occur around the asteroid (21... more 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...
Study of orbital transfers with time constraint and fuel optimization
Journal of Physics: Conference Series, 2017
Neste estudo é analisada a aplicação de um sistema robótico em ambiente espacial, levando em cons... more Neste estudo é analisada a aplicação de um sistema robótico em ambiente espacial, levando em consideração as perturbações causadas à atitude, em virtude do acionamento dos mecanismos robóticos durante a realização de serviços em órbita. O trabalho realizado sugere ganhos consideráveis ao empregar-se modelos que contemplem a correção dinâmica dos erros de posicionamento do referencial do braço robótico, que por sua vez, atua simultaneamente ao controle de atitude do satélite. Atividades robóticas de serviços em órbita têm se tornado comuns devido ao surgimento de novas tecnologias na área de rendezvous, docking e berthing. O braço robótico que serve de objeto de estudo neste trabalho consiste de um manipulador revoluto com três juntas rotativas e três graus de liberdade em configuração Torcional-Rotacional-Rotacional (TRR) movendo-se no espaço. Configuração tal que o confere aplicabilidade diversificada e notória utilidade em atividades de serviço em órbita. A análise dos erros de posicionamento ocasionados pelos movimentos de extensão e rotação do aparato nos possibilita uma visão mais clara sobre as estratégias necessárias para o uso futuro deste tipo de tecnologia. Neste trabalho proponho o uso de um manipulador robótico servindo como ferramenta a contribuir com a efetivação de serviços em ambiente espacial e sanando problemas de controle encontrados quando do uso de outros atuadores para tal finalidade. É, portanto, fundamental que se tenha boa idéia das perturbações causadas à atitude do satélite em decorrência da atuação do braço robótico acoplado. Neste sentido nos concentraremos na análise dos torques perturbadores visto que a base do robô não pode ser considerada, para fins de posicionamento preciso, como sendo de um sistema inercial. A perda da missão pode ser ocasionada por falha prematura de equipamentos do satélite, por isso as agências espaciais têm investido no desenvolvimento de serviços em órbita (OOS -On Orbit Service). OOS inclui diversas atividades de um veículo espacial [1] como montagem, reparo, resgate, aprimoramento, reabastecimento, recuperação e manutenção. Esses serviços podem estender a vida útil dos satélites, melhorar a capacidade dos sistemas espaciais, reduzir custos de operação, e até mesmo contribuir para a mitigação dos detritos espaciais. É imprescindível o domínio das técnicas de rendezvous e docking, bem como, do controle de atitude para executar serviços em órbita em missões do tipo montagem de grandes unidades, reabastecimento de estações, troca de tripulação entre veículos, reparo de satélites, entre outras. Primeiramente, obtemos um modelo que represente, por meio de um algoritmo implementado, a forma como responde um manipulador robótico durante simulação com parâmetros controlados. A configuração de robô articulado (antropomórfico) ou revoluto assemelha-se a um braço humano. Observa-se que é possível calcular a posição cartesiana no espaço, bem como a orientação do punho, com base no conhecimento dos ângulos das juntas. Este equacionamento é conhecido como cinemática direta. O cálculo das posições angulares das juntas a partir da posição no espaço do órgão terminal consiste, portanto, na cinemática inversa. O cálculo da cinemática, tanto direta quanto inversa, requer o conhecimento do comprimento dos elos com precisão adequada, bem como os ângulos de torção entre juntas, ou seja, necessitamos definir tais constantes em nosso modelo computacional e para tanto a chamada notação de Denavit-Hartenberg permite obter o conjunto de equações que descreve a cinemática de uma junta com relação à junta seguinte vide . Assim podemos ter uma visão matematizada da estrutura do robô para inserirmos no algoritmo. A Fig. e a Fig. mostram o mecanismo do qual obtivemos as equações para solução da cinemática inversa, ou seja, dada a posição desejada para o órgão terminal encontram-se os ângulos das juntas capazes de levar a extremidade do robô a tal posição. A modelagem do comportamento do satélite é obtida por meio do Satellite Attitude Simulator (SAS) desenvolvido em [3] e [4], onde o movimento de atitude é calculado a cada passo da simulação. Na arquitetura do simulador Fig. , a atitude de referência é comparada continuamente com a posição angular atual do veículo espacial. Um sinal de erro é gerado por meio da
Resumo Objetiva-se calcular e simular manobras orbitais de um veículo espacial considerando aspe... more Resumo Objetiva-se calcular e simular manobras orbitais de um veículo espacial considerando aspectos construtivos do sistema de propulsão e controle de trajetória em malha fechada. Inicialmente consideram-se propulsores ideais capazes de aplicar empuxo de magnitude infinita, já que a variação de velocidade ocorre instantaneamente. Assim, obtêm-se as manobras impulsivas ótimas, do ponto de vista de consumo de combustível, por meio da solução do problema de Lambert (Two Point Boundary Value Problem). Então, essas manobras são simuladas considerando agora um modelo mais realista do sistema de propulsão. Devido à impossibilidade de aplicação de empuxo infinito a manobra orbital deve ser distribuída em um arco propulsivo em torno do ponto de aplicação determinado pela solução do problema de Lambert. Neste arco propulsivo utiliza-se propulsão contínua com empuxo limitado à máxima capacidade dos propulsores. Entretanto, o efeito do arco propulsivo não é exatamente igual à aplicação de um impulso, esta diferença produz um desvio da órbita final alcançada com relação à órbita de referência. Esse desvio é função da magnitude do impulso desejado, da capacidade do sistema de propulsão e das características do sistema de controle de trajetória utilizado. A avaliação desse desvio é de extrema relevância na análise de missão de veículos espaciais e no dimensionamento do sistema de controle de trajetória. Portanto neste trabalho avaliou-se como a capacidade dos propulsores e a magnitude dos impulsos influenciam nos erros da trajetória quando considerado um modelo mais realista ao invés do caso ideal representado pela abordagem impulsiva. A avaliação dos erros da trajetória torna-se ainda mais relevante se os efeitos das perturbações orbitais que atuam no veículo forem considerados. Assim, os desvios na trajetória devido às perturbações provocadas pelo termo J2 do potencial gravitacional terrestre, pelo arrasto atmosférico, pela pressão de radiação solar, pelo albedo terrestre e pelas atrações gravitacionais do Sol e da Lua foram inseridas na simulação. Palavras-chave Astrodinâmica, satélites artificiais, movimento orbital, manobra orbital.
Space Science Reviews, 2010
Using the Earth albedo model and the orbital dynamics model developed as part of the First Look P... more Using the Earth albedo model and the orbital dynamics model developed as part of the First Look Project (Fast Initial In-Orbit Identification of Scientific Satellites) the terrestrial albedo is evaluated considering the orbits of some scientific missions as Gravity Probe B, MICROSCOPE and STEP. The model of the Earth albedo is based on the reflectivity data measured by NASA's Earth Probe satellite, which is part of the TOMS project (Total Ozone Mapping Spectrometer). The reflectivity data are available daily, on line at the TOMS website, and they fluctuate because of changes in clouds and ice coverage and seasonal changes. The data resolution partitions the Earth surface into a number of cells. The incident irradiance on each cell is used to calculate total radiant flux from the cell. With the radiant flux from each cell, the irradiance at the satellite is calculated.
Journal of Aerospace Technology and Management, 2024
Understanding of various aerodynamic factors involved in flight trajectories is fundamental to de... more Understanding of various aerodynamic factors involved in flight trajectories is fundamental to design launch vehicles. First and foremost, computer simulation is an efficient way of predicting its behavior in the movement across the atmosphere. Considering that the available Brazilian version of Analysis, Simulation and Trajectory Optimization Software for Space Applications (Astos) does not simulate a controlled vehicle in six degrees of freedom (DoF), the aim of this article is to complement the Astos outcomes, particularly evaluating the trajectory of a controlled launch vehicle from liftoff to orbit injection, considering the model of rigid body dynamics with a six DoF. This approach carried out with an in-house developed simulator called Scott that simulated a multistage launcher with three flight configurations. In the Scott computer program, a launcher was modeled with differential equations in six DoF, coupled axes attitude control system, and aerodynamic coefficients that changed as a function of Mach number. These features improved the results generated by Astos software for the same configurations and the same initial conditions. Additionally, the results provided by Scott were close to actual vehicle in terms of attitude change and Mach number reached.
Journal of Aerospace Technology and Management, 2014
One issue the design team has to face in the process of building a new spacecraft, is to define i... more One issue the design team has to face in the process of building a new spacecraft, is to define its mechanical and electrical architecture. The choice of where to place the spacecraft´s electronic equipment is a complex task, since it involves simultaneously many factors, such as the spacecraft´s required position of center of mass, moments of inertia, equipment heat dissipation, integration and servicing issues, among others. Since this is a multidisciplinary task, the early positioning of the spacecraft´s equipment is usually done "manually" by a group of system engineers, heavily based on their experience. It is an interactive process that takes time and hence, as soon as a feasible design is found, it becomes the baseline. This precludes a broader exploration of the design space, which may lead to a suboptimal solution, or worse to a design that will have to be modified later. Recently, it has been shown the potential benefits of automating the process of spacecraft´s equipment layout using optimization techniques. In this paper, a prototype of an Excel ® based tool for multidisciplinary spacecraft equipment layout conception is described. Provided the geometric dimensions, mass and heat dissipation of the equipment, and the available positioning area, the tool can automatically generate many possible trade-off solutions for the layout. It allows the user to set specific equipment to specific areas of positioning, and different combinations of objective functions can be used to drive the design. The features of the tool are shown in a simplified three dimensional problem.
The use of consecutive collision orbits to obtain
56th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law, Oct 17, 2005
Proceeding Series of the Brazilian Society of Computational and Applied Mathematics, Feb 14, 2018
Resumo. Este trabalho tem como principal objetivo analisar a perturbação que cada uma das luas ga... more Resumo. Este trabalho tem como principal objetivo analisar a perturbação que cada uma das luas galileanas Io, Europa, Ganimedes e Calisto, exerce na trajetória de um veículo espacial emórbita ao redor de Júpiter. Além disso, como Júpiter possui várias luas, este estudo considerou a perturbação de 62 luas do planeta, as quais são divididas em 6 grupos: luas galileanas, grupo Almateia, Himalaia, Ananke, Carme e Pasife. Todas estas luas foram consideradas para se ter uma ideia da influência destas perturbações na trajetória do veículo.
Evaluation of the Influence of zonal and sectorial harmonics in the orbit of an lunar satellite
Journal of physics, Oct 1, 2017
This paper presents a study about the tetrahedral layout of four satellites in a way that every h... more This paper presents a study about the tetrahedral layout of four satellites in a way that every half-orbital period this set groups together while flying in formation. The formation is calculated analyzing the problem from a geometrical perspective and disposed by precisely adjusting the orbital parameters of each satellite. The dynamic modelling considers the orbital motion equations. The results are analyzed, compared and discussed. A detection algorithm is used as flag to signal the regular tetrahedron's exact moments of occurrence. To do so, the volume calculated during the simulation is compared to the real volume, based on the initial conditions of the exact moment of formation and respecting a tolerance. This tolerance value is stablished arbitrarily depending on the mission and the formation's geometrical parameters. The simulations will run on a computational environment.
Simulation of the trajectories described by a space vehicle around the asteroid 243 Ida and its natural satellite Dactyl
Journal of physics, Oct 1, 2017
The asteroid 243 Ida located in the asteroid belt, between Mars and Jupiter, is the fourth larges... more 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.
Journal of physics, Oct 1, 2019
Artificial satellites in low Earth orbit have as main disturbance the atmospheric drag, which is ... more 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.
Analysis of the passage of a spacecraft between the Van Allen belts considering a low and high solar activity
Journal of physics, Oct 1, 2017
In this paper, a novel solution for the mixed actuators problem is proposed. This approach uses a... more In this paper, a novel solution for the mixed actuators problem is proposed. This approach uses a multiobjective optimization technique for commanding a group of actuators (thrusters, reaction wheels and magnetic torqrods) simultaneously optimizing a set of objective functions. The multiobjective problem is formulated using combinatorial combinations of the control signal for composing the decision space. The proposed model, called Actuator Multiobjective Command Method (AMCM), is included in a complete guidance, navigation and control loop applied to the final approach rendezvous scenario. Its performance, compared to different actuators configurations, has been evaluated in numerical simulations. In addition, AMCM has been integrated and tested in the hardware-in-the-loop rendezvous and docking simulator of the German Aerospace Center. Results have indicated effectiveness and robustness in both purely numerical and real-time simulations.
Uploads
Papers by evandro marconi rocco