Mission architecture for Mars exploration based on small satellites and planetary drones

Purpose: The purpose of this paper is to deal with the study of an innovative unmanned mission to Mars, which is aimed at acquiring a great amount of detailed data related to both Mars’ atmosphere and surface. Design/methodology/approach: The Mars surface exploration is conceived by means of a fleet of drones flying among a set of reference points (acting also as entry capsules and charging stations) on the surface. The three key enabling technologies of the proposed mission are the use of small satellites (used in constellation with a minimum of three), the use of electric propulsion systems for the interplanetary transfer (to reduce the propellant mass fraction) and lightweight, efficient, drones designed to operate in the harsh Mars environment and with its tiny atmosphere. Findings: The low-thrust Earth-Mars transfer is designed by means of an optimization approach resulting in a duration of slightly more than 27 months with a propellant amount of about 125 kg, which is compatible with the choice of considering a 500 kg-class spacecraft. Four candidate drone configurations have been selected as the result of a sensitivity analysis. Flight endurance, weight and drone size have been considered as the driving design parameters for the selection of the final configuration, which is characterized by six rotors, a total mass of about 6.5 kg and a flight endurance of 28 minutes. In the mission scenario proposed, the drone is assumed to be delivered on the Mars surface by means of a passive entry capsule, which acts also as a docking station and charging base. Such a capsule has been sized both in terms of mass (68 kg) and power (80 W), showing to be compatible with 500 kg-class spacecraft. Research limitations/implications: As a general conclusion, the study shows the mission concept feasibility. Practical implications: The concept would return incomparable scientific data and can be also be potentially implemented with a relatively low budget exploiting of the shelf components to the larger extent, small identical spacecraft buses and modular low-cost drones. Originality/value: The innovative mission architecture proposed in this study aims at providing a complete coverage of the surface and lowest atmospheric layers. The main innovation factor of the proposed mission consists in the adoption of small multi-copter UAVs, also called “drones,” as remote-sensing platforms

Studio preliminare di missioni spaziali con modelli a tre corpi

The study of space missions through models more accurate then a two body model has received a considerable impulse within the scientific community in the last decade. A scheme in which two larger masses determine the motion of a spacecraft which does not modify their gravitational field can be considered satisfactory for the study of a variety of space vehicle trajectories. The possibilities offered by this kind of approach span far beyond the range of the traditional keplerian approach and enable conceiving new types of mission. The present thesis deals with the restricted three body problem, its formulation and its solutions. The different types of trajectories that can be identified by this approach are analysed and the tools that can be used for practical mission implementation are illustrated. In the first part of this work the restricted three body problem is analysed from a theoretical point of view. The steps through which mission design tools can be derived from such a theoretical background are then reviewed. A revised formulation of mathematical results that can be obtained from the theory is presented and the applicability of such results to various missions of potential interest is discussed. Software tools that can be used to describe the theoretically determined space structures, periodic solutions and, in general, for a practical implementation of the theory are described. The second part deals more directly with applications with an inherent three-body character and which could not be designed otherwise. In particular, different possibilities of Earth-Moon transfer, both chemical and hybrid, periodic orbits around the equilibrium points of the three body system and transfer trajectories to and from such orbits are examined in detail

Low Thrust Trajectories in Multi Body Regimes

More and more stringent and unique mission requirements motivate to exploring solutions, already in the preliminary mission analysis phase, going far beyond the classical chemical-Keplerian approach. The present dissertation deals with the analysis and the design of highly non linear orbits arising both from the inclusion of different gravitational sources in the dynamical models, and from the use of electric system for primary propulsion purposes. The equilibrium of different gravitational fields, on one hand, permits unique transfer solutions and operational orbits, on the other hand, the high thrust efficiency, characteristic of an electric device, reduces the propellant mass required to accomplish the transfer. Each of these models, and even better their combination, enables trajectories able to satisfy mission requirements not otherwise met, first of all to reduce the propellant mass fraction of a given mission. The inclusion of trajectory arcs powered by an electric thruster, providing a low thrust for extended duration, makes essential the use of optimal control theory in order to govern the thrust law and thus design the required transfers so as to minimizing/maximizing specific indexes. The goal is, firstly, to review the possible advantages and the main limits of dynamical models and, afterward, to define methodologies to preliminary design non-Keplerian missions both in interplanetary contexts and in the Earth-Moon system. Special emphasis is given to the study of dynamical systems through which the main features of the Circular Restricted Three Body Model (the first one among the non-Keplerian models) can be identified, implemented and used. Purely ballistic solutions enabled by this model are first independently explored and after considered as target orbits for electric thrusting phases. Electric powered arcs are used to link ballistic phases arising from the balancing of different gravitational influences. This concept is applied both for the exploration of planetary regions and for interplanetary transfer purposes. Together with low thrust missions to selenocentric orbits designed taking into account both the Earth and the Moon gravity, also transfer solutions toward periodic orbits moving in the Earth-Moon region are presented. These are designed considering electric thrusting arcs and ballistic segments exploring for free specific space regions. In brief, theoretical models deriving from dynamical system theory and from optimal control theory are employed to design non conventional orbits in non linear astrodynamics models

Development and Test of Low Cost Solar Panel Technologies for Small Satellites

This paper presents the activities carried out in collaboration between the University of Pisa and Alta SpA for the development, testing and integration of an efficient, yet inexpensive photovoltaic panel for microsatellite applications. The approach adopted, aimed at reducing cost and developing “low tech” techniques to assembly and qualify solar panels for small satellite applications, uses a printed circuit board designed to optimize the use of external surfaces partially occupied for power generation, where bare cells are installed by means of a double-sided insulating adhesive tape and each cell is covered with cerium doped borosilicate glass, using a controlled volatility silicone. Bonding was performed with a dedicated vacuum bag technique, developed in-house. This method achieves a significant cost reduction with respect to traditional techniques, while retaining high performance and reliable repeatability and avoiding complex technological procedures during the integration. A prototype solar panel was manufactured, tested and integrated on the UniSat-5 small spacecraft by GAUSS Srl in preparation of a flight scheduled for late 2013. Thorough mechanical testing was performed as a part of the integration with UniSat-5. The panels manufactured during the development programme were subject to electrical characterization to evaluate the current-voltage characteristic curve and the efficiency of the array and to thermal vacuum tests according to ECSS standards to estimate the outgassing properties of the protoflight model. For both tests, a low cost experimental setup was developed on purpose. The recorded flight unit total mass loss (TML) is well under the acceptable limits, so that the panel was accepted for space flight. In-orbit validation of the panel is expected with the upcoming flight of UniSat-5. The techniques and procedures developed under this programme allow for quick and inexpensive manufacture of reliable solar arrays, specially suited for micro- and nano-satellites. To improve the thermal and mechanical properties of the solar array, a substrate in carbon fibre composite laminate is under investigation. A thermal analysis is developed to characterize and compare the thermal response of the solar array with different substrate subjected to space heat flux

Slit FEEP Thruster Performance with Ionic Liquid Propellant

By replacing the liquid metal propellant with a ionic liquid, it is possible to develop a new, simplified FEEP system that combines most of the heritage and the advantages of the linear slit geometry with the easy of handling and operation of a more benign propellant. In view of the development of such Ionic Liquid FEEP thruster (IL-FEEP), an internal development activity is underway at Alta, aimed at the design and testing of an innovative linear slit thruster derived from the cesium experience. This paper presents the results of recent experimental campaigns aimed at assessing the performance of linear slit FEEP emitters fed with a ionic liquid propellant. For the first time, beam composition was evaluated using a time-of-flight mass spectrometry technique, allowing for a reliable estimate of the thruster’s specific impulse

Design Criteria of Remote Sensing Constellations of Small Satellites with Low Power Electric Propulsion and Distributed Payloads

The recent explosion in proposed microsatellite missions is based on the possibility to mass-produce cheap platforms capable to deliver acceptable performance over a limited lifetime. The assumption behind such scheme is that individual microsatellites are expected/allowed to fail in reasonable numbers, the resulting degradation of constellation performance being limited due to the large population of active spacecraft. We argue that cheap platforms do not necessarily need to be seen as disposable assets, so that low cost constellations featuring a low number of microsatellites may nevertheless be capable of remarkable performance. The key technology needed to enable such feat is low power electric propulsion, whereby microsatellites are allowed to acquire and maintain precisely tuned orbital locations, compensate atmospheric drag to fly longer, and de-orbit safely at end of life. A number of such microsatellites may be fitted with an instrument each from a suite of different sensors operating in various spectral bands. The constellation would operate as an actively controlled system, with the individual instruments providing well coordinated raw data that may be processed using data fusion techniques to yield the final product. Starting from the proven performance of a currently available low power Hall thruster, we present general design criteria for constellations based on a 50 kg-class microsatellite bus. The potential benefits of such technology are outlined with respect to applications such as precision farming, urban area monitoring, and dual use land surveillance

Recent Developments in Ionic Liquid Field Emission Electric Propulsion

The study herein reported was aimed at the characterization of the plume of a ionic liquid fed, linear slit FEEP thruster, in terms of composition and velocity of the constituents. Ionic liquid propellants are actively investigated as promising alternatives to alkali metals in field emission thrusters, in order to reduce system cost and ground operation complexity. To this end, a large number of tests was carried out using the EMI-BF4 ionic liquid as a propellant. The thruster was fired in either positive polarity or negative polarities to check the capability to extract anions and cations alone. Then, most of the testing was carried out in alternate polarity mode, in order to avoid electrochemical poisoning of the propellant, due to the unbalanced extraction of charged particles[1]. Such operating mode is believed to be the most promising candidate for flight operation, as it would allow to get rid of an external neutralizer to maintain electrical neutrality of the spacecraft. Ion beam composition was investigated by means of a time-of-flight mass spectrometry technique. The measurements show that the emitted beam is mostly composed of monomers (BF4)-, dimers (C6H11BF4N2) (BF4)- and polymers (C6H11BF4N2)n (BF4)- (with n a function of applied extraction voltage). Under the assumption of a certain beam composition, propellant consumption was indirectly evaluated by means of time integration of the emitted current and independently verified by means of direct observation of the depletion of the propellant reservoir. The estimated resulting specific impulse is around 1400 s. The thruster behaviour resulted quite variable, especially when operated at high voltage levels in continuous polarity mode. Better performance was registered in alternate polarity operation with an alternation period of several tens of a seconds at extracted current of just a few µA

Low-Thrust Propulsion Systems for Small Satellites

Small platforms represent valuable options for small scientific and Earth observation missions. One of the main challenges for microspacecrafts with launch mass below approx. 100 kg is the inclusion of a propulsive subsystem. Such addition would significantly enhance the performance of these platforms, broadening the possible applications and/or extending the operational lifetime. For these applications, electric propulsion systems are more suitable than classical chemical systems as they allow a larger payload mass fraction reducing the propellant mass requirement. The aim of this study is to investigate about state of the art of electric propulsion options for small spacecraft. A selection of possible electric propulsion systems for small satellite based on several requirements (power, thrust, specific impulse) is presented. This paper discusses four typical orbital manoeuvres of fundamental relevance for satellites in low Earth orbit: scenario No. 1 considers a 350 km decrease of orbital altitude; scenario No. 2 is about drag compensation for 90 days at very low altitude. Scenario No. 3 is a combined manoeuvre to change both semi major axis (by 150 km) and inclination (by 0.563 deg) at once. The fourth scenario is about orbit circularization. Results are thus normalized to obtain dimensionless parameters to be compared. Our analysis shows that an electric propulsion system offers significant advantages for small satellites in low Earth orbits in case of altitude variation, moderate inclination change, circularization and drag compensation at very low altitude

Optimal low-thrust transfers between libration point orbits

Over the past three decades, ballistic and impulsive trajectories between libration point orbits (LPOs) in the Sun–Earth–Moon system have been investigated to a large extent. It is known that coupling invariant manifolds of LPOs of two different circular restricted three-body problems (i.e., the Sun–Earth and the Earth–Moon systems) can lead to significant mass savings in specific transfers, such as from a low Earth orbit to the Moon’s vicinity. Previous investigations on this issuemainly considered the use of impulsive maneuvers along the trajectory. Here we investigate the dynamical effects of replacing impulsive V’s with low-thrust trajectory arcs to connect LPOs using invariant manifold dynamics. Our investigation shows that the use of low-thrust propulsion in a particular phase of the transfer and the adoption of a more realistic Sun–Earth–Moon four-body model can provide better and more propellant-efficient solution. For this purpose, methods have been developed to compute the invariant tori and their manifolds in this dynamical modelPeer ReviewedPostprint (published version

Performance of Fatty Liver Index in Identifying Non-Alcoholic Fatty Liver Disease in Population Studies. A Meta-Analysis

Background. Fatty liver index (FLI) is a non-invasive tool used to stratify the risk of non-alcoholic fatty liver disease (NAFLD) in population studies; whether it can be used to exclude or diagnose this disorder is unclear. We conducted a meta-analysis to assess the prevalence of NAFLD in each FLI class and the performance of FLI in detecting NAFLD. Methods. Four databases were searched until January 2021 (CRD42021231367). Original articles included were those reporting the performance of FLI and adopting ultrasound, computed tomography, or magnetic resonance as a reference standard. The numbers of subjects with NAFLD in FLI classes <30, 30–60, and 60, and the numbers of subjects classified as true/false positive/negative when adopting 30 and 60 as cut-offs were extracted. A random-effects model was used for pooling data. Results. Ten studies were included, evaluating 27,221 subjects without secondary causes of fatty liver disease. The prevalence of NAFLD in the three FLI classes was 14%, 42%, and 67%. Sensitivity, specificity, positive predictive value, negative predictive value, likelihood ratio for positive results, likelihood ratio for negative results, and diagnostic odds ratio were 81%, 65%, 53%, 84%, 2.3, 0.3, and 7.8 for the lower cut-off and 44%, 90%, 67%, 76%, 4.3, 0.6, and 7.3 for the higher cut-off, respectively. A similar performance was generally found in studies adopting ultrasound versus other imaging modalities. Conclusions. FLI showed an adequate performance in stratifying the risk of NAFLD. However, it showed only weak evidence of a discriminatory performance in excluding or diagnosing this disorder