Systems design of a hybrid sail pole-sitter

This paper presents the preliminary systems design of a pole-sitter. This is a spacecraft that hovers over an Earth pole, creating a platform for full hemispheric observation of the polar regions, as well as direct-link telecommunications. To provide the necessary thrust, a hybrid propulsion system combines a solar sail with a more mature solar electric propulsion (SEP) thruster. Previous work by the authors showed that the combination of the two allows lower propellant mass fractions, at the cost of increased system complexity. This paper compares the pure SEP spacecraft with the hybrid spacecraft in terms of the launch mass necessary to deliver a certain payload for a given mission duration. A mass budget is proposed, and the conditions investigated under which the hybrid sail saves on the initial spacecraft initial mass. It is found that the hybrid spacecraft with near- to mid-term sail technology has a lower initial mass than the SEP case if the mission duration is 7 years or more, with greater benefits for longer duration missions. The hybrid spacecraft with far-term sail technology outperforms the pure SEP case even for short missions

PROGETTAZIONE DI UN PROTOTIPO DI ENDOREATTORE AUTO PRESSURIZZATO A PEROSSIDO DI IDROGENO ED ETANO

Negli ultimi anni il costo associato alle varie operazioni a terra riguardanti propellenti non criogenici attualmente in uso come l’idrazina N2H4 ed i suoi derivati (MMH, UDMH) e il tetrossido di azoto (NTO) è aumentato notevolmente. La graduale presa di coscienza dell’elevata tossicità di tali propellenti ha prodotto negli anni una progressiva diminuzione dei limiti di esposizione previsti e quindi un conseguente aumento dei costi inerenti a tutte le procedure di manipolazione e stoccaggio e riguardanti la sicurezza e l’incolumità del personale che se ne occupa. Questo ha portato ad un aumento di interesse verso valide alternative che presentino una minore tossicità, specialmente per applicazioni su satelliti di piccole e medie dimensioni, dato il maggior peso dei costi a terra rispetto al costo dell’intera missione. Per tali applicazioni il perossido d’idrogeno (HP, H2O2) è uno dei più promettenti propellenti stoccabili “verdi” in virtù del suo costo contenuto. Ha un ridotto impatto ambientale ed è caratterizzato da una bassa tossicità e da una certa facilità nelle procedure di produzione, manipolazione e stoccaggio. Al momento Alta S.p.A. sta lavorando allo sviluppo di letti catalitici per la decomposizione del perossido d’idrogeno e alla loro applicazione in endoreattori monopropellenti. L’esperienza acquisita in questo settore dà la possibilità di analizzare le prestazioni propulsive di un innovativo motore bi- propellente auto-pressurizzato a perossido d’idrogeno ed etano (C2H6) secondo il concetto FVP (Fuel vapor Pressurization), nel quale si sfrutta l’alta pressione di vapore di un idrocarburo leggero per la pressurizzazione di entrambi i propellenti contenuti in un singolo serbatoio separati da un diaframma mobile (doppio serbatoio). Questo sistema di pressurizzazione promette una significativa riduzione dei costi, del numero degli elementi e della massa totale del sistema. La presente tesi tratta la progettazione di un prototipo di endoreattore da 50 N di spinta per la sperimentazione a terra del concetto FVP applicato a propellenti non tossici. Dopo una breve introduzione sulle prestazioni dei motori a perossido d’idrogeno, nel Capitolo 2 sono descritte le proprietà fisiche dei due propellenti. Nel capitolo 3 è illustrata l’analisi termodinamica del doppio serbatoio e la verifica dell’equivalenza ad una configurazione a serbatoi separati più adeguata per una sperimentazione a terra. Nel Capitolo 4 è affrontato il dimensionamento del propulsore comprendente camera di spinta (camera di combustione e ugello di spinta), camera di miscelamento dei propellenti, e letto catalitico. Il Capitolo 5 mostra il dimensionamento delle linee di alimentazione dei propellenti. Il Capitolo 6 tratta il sistema di raffreddamento della camera di spinta. Il lavoro di tesi si è svolto presso Alta S.p.A.,sotto la supervisione del Prof. Luca d’Agostino, e degli Ing. Lucio Torre, Angelo Pasini e Luca Romeo

Fast E-sail Uranus entry probe mission

The electric solar wind sail is a novel propellantless space propulsion concept. According to numerical estimates, the electric solar wind sail can produce a large total impulse per propulsion system mass. Here we consider using a 0.5 N electric solar wind sail for boosting a 550 kg spacecraft to Uranus in less than 6 years. The spacecraft is a stack consisting of the electric solar wind sail module which is jettisoned roughly at Saturn distance, a carrier module and a probe for Uranus atmospheric entry. The carrier module has a chemical propulsion ability for orbital corrections and it uses its antenna for picking up the probe's data transmission and later relaying it to Earth. The scientific output of the mission is similar to what the Galileo Probe did at Jupiter. Measurements of the chemical and isotope composition of the Uranian atmosphere can give key constraints to different formation theories of the Solar System. A similar method could also be applied to other giant planets and Titan by using a fleet of more or less identical probes.Comment: 14 pages, 5 figures, Meudon Uranus workshop (Sept 16-18, 2013) special issue of Planetary and Space Scienc

Constrained Large Angle Reorientation Manoeuvres of a Space Telescope Using Potential Functions and a Variable Control Gain

In this paper we study large angle rotational maneuvers of a space telescope with pointing constraints. The spacecraft attitude control design is formulated and solved by means of potential functions, thus simplifying the problem of frequent reorientation maneuvers. A novel approach is proposed, where a time varying control gain is chosen such that its instantaneous value depends both on the spacecraft kinetic energy and on the distance of the spacecraft from the forbidden directions. As a result, the spacecraft is able to reach points in the potential field arbitrarily close to a constraint and to maneuver with autonomous capability of guidance and control. A case study illustrates the effectiveness of the proposed methodology

Optimal Nodal Flyby with Near-Earth Asteroids Using Electric Sail

The aim of this paper is to quantify the performance of an Electric Solar Wind Sail for accomplishing flyby missions toward one of the two orbital nodes of a near-Earth asteroid. Assuming a simplified, two-dimensional mission scenario, a preliminary mission analysis has been conducted involving the whole known population of those asteroids at the beginning of the 2013 year. The analysis of each mission scenario has been performed within an optimal framework, by calculating the minimum-time trajectory required to reach each orbital node of the target asteroid. A considerable amount of simulation data have been collected, using the spacecraft characteristic acceleration as a parameter to quantify the Electric Solar Wind Sail propulsive performance. The minimum time trajectory exhibits a different structure, which may or may not include a solar wind assist maneuver, depending both on the Sun-node distance and the value of the spacecraft characteristic acceleration. Simulations show that over 60% of near-Earth asteroids can be reached with a total mission time less than 100 days, whereas the entire population can be reached in less than 10 months with a spacecraft characteristic acceleration of 1 mm/s(2)

Minimum-Time Trajectories of Electric Sail with Advanced Thrust Model

The Electric Solar Wind Sail is an advanced propulsion system concept that, similar to the more conventional solar sail, is able to generate a propulsive thrust without any propellant. The main performances of such a propulsion system have been studied in different mission scenarios and are reported in the literature. However, the analyses available so far are based on a simplified thrust model that neglects the effect of the spacecraft attitude on both the thrust modulus and its direction. The recent availability of a refined thrust model requires a critical reappraisal of the simulation results and a new analysis of the optimal trajectories of a spacecraft equipped with such a propulsion system. The aim of this paper is to review the different thrust models used over the last years for mission analysis purposes, and to illustrate the optimal control law and the corresponding minimum-time trajectories that can be obtained with the new, refined, thrust model. The study highlights new analytical relations for the propulsive thrust as a function of the spacecraft attitude, whereas simple and accurate closed-form equations are also proposed for the study of a classical circle-to-circle coplanar heliocentric orbit transfer

Analysis of spacecraft motion under constant circumferential propulsive acceleration

This paper reassesses the classical circumferential-thrust problem, in which a spacecraft orbiting around a primary body is subjected to a propulsive acceleration of constant modulus, whose direction is in the plane of the parking orbit and orthogonal to the spacecraft-primary line. In particular, a new formulation is proposed to obtain a reduction in the number of differential equations required for the study of the spacecraft propelled trajectory. The mathematical complexity of the problem may be further reduced assuming that both the propulsive acceleration modulus and the spacecraft distance from the primary body are sufficiently small. In that case, an approximate model is able to accurately describe the characteristics of the propelled trajectory when the parking orbit is circular. Finally, using the data obtained by numerical simulations, the approximate model is extended to generate a set of semi-analytical equations for the analysis of a classical escape mission scenario

Heliocentric Trajectory Analysis of Sun-pointing Smart Dust with Electrochromic Control

A smart dust is a micro spacecraft, with a characteristic side length on the order of a few millimeters, whose surface is coated with electrochromic material. Its orbital dynamics is controlled by exploiting the differential force due to the solar radiation pressure, which is obtained by modulating the reflectivity coefficient of the electrochromic material within a range of admissible values. A significant thrust level can be reached due to the high values of area-to-mass ratio of such a spacecraft configuration. Assuming that the smart dust is designed to achieve a passive Sun-pointing attitude, the propulsive acceleration due to the solar radiation pressure lies along the Sun-spacecraft direction. The aim of this paper is to study the smart dust heliocentric dynamics in order to find a closed form, analytical solution of its trajectory when the reflectivity coefficient of the electrochromic material can assume two values only. The problem is addressed by introducing a suitable transformation that regularizes the spacecraft motion and translates the smart-dust dynamics into that of a linear harmonic oscillator with unitary frequency, whose forcing input is a boxcar function. The solution is found using the Laplace transform method, and afterwards the problem is generalized by accounting for the degradation of the electrochromic material due to its exposition to the solar radiation. Three spacecraft configurations, corresponding to low, medium and high performance smart dusts, are finally used to quantify the potentialities of these advanced devices in an interplanetary mission scenario