Nanosatelliti – Tesi Svolte – New

In recent years, many industries and research institutes are trying to access the space. For this, rockets are always used to bring the satellite in the orbit of earth. This makes a lot of effort and tedious that leads to make a small satellite with same pitcher that easily launches in the space along with large number of other small satellites simultaneously. Their launching cost becomes economical when shared by different manufacturer and also affordable for universities and small companies. Secondly, the small satellites are going to decrease in weight, dimension and cost day by day but it increase the complexity of the system progressively. A continuing miniaturization of electronic components has played a major role to decrease the complexity of the system and create an innovation project that no one could ever think about it.
The innovative project known as ‘AraMiS’ satellite is followed after many experiments and hard working done by many researchers. It is basically a nanosatellite whose weight is between 1 to 10kg. This design gives a low cost and high performance approach to the new world. The reason for high performance is that it has power management subsystem that attains the maximum solar power generated by solar cells. Different numbers of solar cells are used according to its application that should also meet the requirement of budget.
Every satellite system must ensure the critical functions. In particular:
 Power Management System
 Position Control System
 Housekeeping Sensor
 Management and Analysis of Data Control in the satellite
 Telecommunication System
In addition, the components of the satellite are affected by different noise frequency that is generated by internal and external sources. For the external noise, the structure is completely metallic body and it has a good shielding against electromagnetic emissions (EMI). For the internal noise, it creates interference between the various boards or within a same board. It is controllable when placing the appropriate positioning of ground planes of both Analog and Digital units.
Before the end of the project, all the components of small satellite must be described in Visual Paradigm software. It makes for easily understanding and rapidity to define projects in a clear and efficient way. UML is used for the design and documentation of the AraMiS project. The functionality of each and every module of the project must be mentioned in UML.

Un sistema aerospaziale è un qualsiasi sistema che vive nell’aria,
entro e oltre l’atmosfera, come ad esempio satelliti, sonde spaziali, aerei,
palloni sonda, eccetera.
Con il termine “satellite” si indicano tutti gli oggetti orbitanti attorno
ad un corpo celeste che sono di dimensioni molto minori rispetto ad un
pianeta. Si possono distinguere in
− satelliti naturali, cioè un qualunque corpo celeste che orbiti
attorno ad un corpo diverso da una stella
Capitolo 1 – Introduzione
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− satelliti artificiali, ovvero tutti gli oggetti orbitanti intorno ad un
corpo celeste che sono stati posti volutamente in quell’orbita
con mezzi tecnologici
Figura 1.1: Il primo satellite artificiale in orbita della storia: lo Sputnik
Le sonde spaziali sono oggetti simili ai satelliti artificiali, ma non
orbitano attorno ad un corpo celeste bensì viaggiano verso pianeti
esterni al sistema solare e sono principalmente veicoli esplorativi.
I satelliti artificiali si possono suddividere in:
− satelliti per le telecomunicazioni, come i satelliti
COSPAS/SARSAT, che spesso sono posizionati in un’orbita
geostazionaria,
− satelliti meteorologici, posizionati sia in orbita geostazionaria
(es. METEOSAT) sia in orbita polare (es. satelliti NOAA)
− satelliti scientifici, ad esempio il Telescopio Spaziale Hubble,
Envisat, Landsat e RapidEye
− satelliti militari, i quali possono avere sia scopi difensivi che
offensivi (ad esempio la rete di satelliti di monitoraggio nucleare
Vela e Geosat)
Capitolo 1 – Introduzione
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− stazioni orbitanti come la Stazione Spaziale Internazionale,
Skylab, Mir.
Per i satelliti artificiali, così come per quelli naturali, è possibile
calcolare la loro velocità di spostamento in orbita. Per fare ciò è
necessario ipotizzare che
− l’orbita del satellite sia circolare
− il satellite si muova attorno ad un corpo puntiforme con una certa
massa
− il satellite sia anch’esso un corpo puntiforme
Un satellite che ruota attorno alla Terra è soggetto alla forza di gravità,
che cerca di farlo precipitare sulla Terra, e alla forza centrifuga, che
cerca di farlo allontanare. Per la legge del moto circolare uniforme, è
noto che la forza centrifuga è data da…

La presente trattazione si inserisce all’interno di una collaborazione con il Dipartimento di Elettronica del Politecnico di Torino per l’approfondimento strutturale del comportamento di particolari elementi di impiego aerospaziale. Tali elementi di tipo modulare presentano particolari vantaggi per la realizzazione di componenti più complessi e multifunzionali che quindi richiedono una caratterizzazione più di dettaglio anche dal punto di vista dinamico-strutturale.
L’impiego di questi elementi permetterebbe l’abbattimento dei costi nella realizzazione di sistemi/componenti aerospaziali di scala ridotta grazie ad un’impostazione di tipo modulare basata sull’idea di un componente multifunzionale. Si tratta di un elemento con il compito tanto di assolvere a funzioni prettamente strutturali quanto di fungere da piattaforma per ospitare i principali sottosistemi che comunemente trovano impiego sulla classe di configurazioni aerospaziali interessata. Trattandosi inoltre di strutture modulari, offrono la possibilità di essere assemblate in geometrie particolarmente vantaggiose, creando configurazioni adatte a svolgere funzioni anche totalmente differenti fra loro (dalla struttura per pannello solare a quella per ospitare telescopi ottici a quella, cubica, adatta per accogliere un generico payload).
Assume quindi un interesse concreto l’analisi strutturale dinamica della piastra in questione, per valutarne frequenze di vibrazione proprie e deformazioni modali nelle varie configurazioni di vincolo e di carico. Inoltre risulta importante definire le metodologie di simulazione il numero di elementi, l’effetto della tipologia di elemento utilizzato e la simulazione dei vincoli. La tesi quindi cerca di
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definire in maniera operativa questi aspetti avendo come obiettivo quello di caratterizzare l’elemento dal punto di vista delle frequenze proprie di vibrazione.
Ci si è avvalsi, per tale scopo, del metodo di calcolo FEM tramite il codice di calcolo Nastran e dello strumento di pre/post processing Patran.
L’approccio seguito per affrontare questo problema prevede di creare un modello agli elementi finiti della piastra procedendo per gradi. Si è partiti dal massimo livello di esemplificazione possibile fino a giungere ad un modello che rappresentasse la struttura “reale”, arricchendo e modificando opportunamente quello creato in precedenza.
Gli elementi di complicazione presenti nel componente studiato sono dovuti principalmente alla natura non esclusivamente strutturale dello stesso; essendo infatti predisposto ad ospitare svariati elementi circuitali, esso è composto da più strati di geometrie e materiali non uniformi. Si è trattato inoltre il caso particolarmente interessante, studiato nel Capitolo 9, della struttura collegata nel proprio centro ad una massa costituita da un insieme motore-ruota di inerzia. Questa rappresenta, fra le configurazioni standard previste per la piastra strumentata, quella di massima sollecitazione.
Il lavoro svolto si è infine concluso con la realizzazione di un modello semplificato di una configurazione realistica a geometria cubica contenente la piastra analizzata in precedenza.
Per ogni stadio dell’analisi si sono trattate le condizioni al contorno più frequenti e si è cercato, laddove possibile, di confrontare i dati ottenuti dal calcolatore con quelli calcolati avvalendosi delle teorie descrittive di volta in volta più adatte. Si sono quindi commentati i risultati più importanti.

Il presente lavoro è frutto di una collaborazione con il dipartimento di elettronica e telecomunicazioni ed ha lo scopo di analizzare il comportamento modale e termico di alcuni componenti multifunzionali presenti sul modulo satellitare Aramis, progetto sviluppato all’interno dell’ateneo sotto la guida del professor Reyneri.
Il progetto Aramis si riferisce alla realizzazione di un satellite low cost di tipo cubesat, caratterizzato dall’assemblaggio di componenti modulari multifunzionali pre-assemblati e pre-testati, che uniti fra loro da elementi strutturali di giunzione vanno a definire la forma parallelepipeda del satellite.
I “cubesat” sono satelliti cubici di piccole dimensioni ideati sul finire degli anni ’90 nell’università americana Cal.Tec. con scopo di fornire una soluzione standard per il design di picosatelliti, in modo da ridurre costi e tempi di sviluppo e aumentare l’accessibilità alle missioni satellitari e la sostenibilità alla frequenza dei lanci.
Il progetto ha suscitato fin da subito grande interesse, sia in ambiente accademico che aziendale, tant’è che attualmente oltre 100 istituti nel mondo fra università, scuole ed imprese, sono coinvolti nel progetto cubesat col fine di sviluppare soluzioni strutturali sempre più efficienti e payload utili ai vari scopi specifici, con la speranza di poter vedere il loro cubesat in orbita.
Il progetto Aramis, iniziato nel 2007, ha come scopo principale quello didattico e diversi scopi secondari.
Esso viene proposto agli studenti che vogliono cimentarsi nella realizzazione di un progetto spaziale per apprendere quale siano i processi che portano al compimento di una missione satellitare, capire come applicare al meglio le soluzioni tecnologiche disponibili, far nascere nuove collaborazioni tra studenti, aziende e dipartimenti all’interno dell’ateneo, e sollecitare la ricerca di nuove soluzioni e tecnologie utili all’uopo. In più la possibilità di collaborazione al progetto è estesa anche a livello scolastico, in quanto anche diversi licei ed istituti secondari sono stati coinvolti nello sviluppo di un ulteriore payload da inserire sul satellite.
Oltre agli scopi didattici, la missione si pone il fine di far nascere collaborazioni fra l’ateneo ed il mondo delle aziende, rendere funzionante ed operativa la ground station presente nell’ateneo e testare il comportamento alle radiazioni di alcuni componenti hardware da poco presenti sul mercato, i quali andranno a definire il payload del satellite.
1.1 Aramis
Aramis presenta un’architettura innovativa inizialmente pensata per ridurre l’ingombro di satelliti di grosse dimensioni; esso infatti prevede il montaggio dei componenti hardware direttamente sui pannelli che ne definiscono la struttura, i quali vengono denominati “tiles”.
Ogni tile rappresenta una delle sei superfici multi funzionali che definiscono la forma parallepipeda del satellite. La caratteristica multifunzionale di queste superfici deriva dal fatto che ognuna, oltre ad assolvere la funzione strutturale, ha montati su di essa le schede circuitali ed i componenti hardware che di solito vengono montati nel volume interno del satellite, e quindi separatamente dagli elementi che ne definiscono la forma e le proprietà strutturali.
In questo modo si riesce a contenere in maniera molto efficace gli ingombri derivanti dalle schede hardware ed a ottenere un significativo volume vuoto al centro del satellite utilizzabile per altri scopi.
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Figura 1- spaccato di Aramis
Aramis è costruito con quattro tiles laterali rettangolari di dimensione 330 x 165 mm destinate ad ospitare esternamente i pannelli solari ed internamente il payload, il power management system, il sistema di controllo d’assetto e i computer di bordo, e due tiles superiori quadrate di dimensione 165 x 165 mm destinate al sistema di telecomunicazione e scambio dati tramite un antenna UHF posta sulla superficie esterna, una SHF interna ed altri componenti hardware.
Le tiles sono collegate tra loro da correnti profilati a L, attraverso i quali passano le viti di giunzione.
Figura 2- vista frontale PCB esterno della tile con supporti di montaggio celle
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ll payload del satellite consiste in un set di schede hardware, sviluppate per testare in ambiente spaziale diversi microcontrollori da poco sul mercato e di alto interesse tecnologico per le future missioni spaziali.
Il test di questi componenti verrà effettuato inizialmente dal computer system di Aramis, il quale processerà continuamente i dati sul funzionamento di questi dispositivi e li invierà a terra, e successivamente dalla ground station, che ne analizzerà il comportamento in relazione al livello di radiazioni a cui è sottoposto il satellite. Le radiazioni sono causate essenzialmente da ioni pesanti e particelle cariche ad alta velocità, il cui livello d’intensità è misurato da alcuni componenti montati a bordo, nell’unità di misura SEU.
Dato lo scopo della missione, qualsiasi orbita LEO è adeguata, basta che vi sia sufficiente visibilità dalla ground station posta in Torino; questo significa che si dovranno evitare orbite sub-equatoriali con angolo d’inclinazione minore di 45°.
Il presente lavoro ha lo scopo di analizzare il comportamento termo-dinamico di una di queste tiles considerata in tre configurazioni strutturali diverse, e di valutare quale di queste soluzioni sia migliore o più efficiente i termini di massa, rigidezza, e di isolamento termico dei componenti elettronici.
Per ognuna di queste tre configurazioni si determineranno le frequenze naturali di vibrazione trasversale della tile sia in modo analitico, attraverso diversi modelli ingegneristici approssimati, sia in modo simulativo tramite elementi finiti con implementazione in Nastran. Con lo stesso modello agli elementi finiti si determinerà il massimo spostamento trasversale della tile simulando la fase di lancio.
Successivamente si andrà a condurre l’analisi termica della tile per determinarne i differenziali di temperatura caratteristici in determinate condizioni di irraggiamento. Infine si valuterà quale configurazione della tile sia più efficiente in termini di rigidezza strutturale, proprietà inerziale ed isolamento termico.

Lo sviluppo di applicazioni spaziali comporta costi non indifferenti, sia da un punto di vista
puramente economico sia dal lato progettuale. Inoltre si incontrano problemi particolari come
quello dell’assenza di gravità che comportano evoluzioni particolari le quali vanno studiate
appositamente caso per caso. La nascita di studi a costo ridotto come quello a cui si riferisce la
seguente trattazione comporta necessariamente una spinta verso un approccio modulare al quale
bene risponde l’architettura AraMiS.
Dopo una iniziale introduzione a quelle che sono le effettive applicazioni ed a quello a cui
servono, vengono introdotti i termini necessari allo studio di analisi termica per satelliti.
Dopodichè lo studio si divide in tre parti principali ovvero l’analisi strutturale e dunque le teorie
che permettono lo studio delle autovibrazioni della tile in oggetto, importante punto di
valutazione della struttura, l’analisi termica vera e propria ed infine gli esperimenti in laboratorio
che ne caratterizzano e verificano i risultati.
Buona parte è dedicata alla modellizzazione di una particolare struttura che grazie alle sue
proprietà in campo sia strutturale che termico viene sempre più utilizzata, in particolare in
ambiente aerospaziale. Si tratta della struttura a sandwich in configurazione con core in
honeycomb e ne verranno presentati diversi modelli equivalenti a seconda dello studio che
dobbiamo fare. Ad esempio verrà omogeneizzato nel caso strutturale e semplificato nel caso
termico. Di particolare importanza lo studio basato sui fattori di vista per la conduzione attraverso
l’honeycomb che permette di considerare l’apporto dovuto all’irraggiamento.
I test effettuati nei laboratori di microelettronica del Dipartimento di Elettronica e
Telecomunicazioni del Politecnico di Torino, verranno messi a confronto con i valori teorizzati
dando risultati decisamente concordanti con le misure verificando sia i calcoli che i modelli
ipotizzati.

This work has been done by Sevil M. Sadigh under supervision of Professor Leonardo Reyneri at the Polytechnic University of Turin1, Italy. Sevil M. Sadigh is PhD student from Iran and she is doing her PhD thesis, at the moment. A part of her thesis is about satellite attitude control. She has worked on this part of her PhD thesis at the Department of Electronics and Telecommunications, as a visiting research scholar for a period of six months, from May 5th to Nov 5th 2017.
The works done in this period is described in this text as follows. In section 1, coordinate frames such as inertia, body, and orbit are described to determine the satellite attitude. Some types of orbital classifications like as centric classifications, altitude classifications for geocentric orbits, eccentricity classifications, and inclination classifications are described in section 2. Also, some necessary information like as orbital elements, orbital period and earth magnetic field are explained. In section 3, satellite attitude and the types of the rotation are described. Also, how to convert them to each other are explained. Then satellite kinematic and dynamic equations are described in section 4. Attitude actuators, their characteristics, advantages and disadvantages and their operations are explained in section 5. Attitude control is described in section 6. Attitude control approach used in this work is SMC2. SMC method is explained its advantages and disadvantages is described then the controller is designed for satellite attitude control with magnetorquer and reaction wheel actuators, in this section. In section 7, the proposed attitude control is simulated for the nanosatellites with several combinations of the actuators. At first, a sample attribute motion for tracking has been described. Then, attitude dynamic and kinematic of the satellite has been explained. The numerical model for the attitude control has been described in the next section. Finally, the performed simulations are explained. In the final section, the conclusion and future works are described.

PiCPoT è il nome del primo nanosatellite progettato dal Politecnico di Torino. La
sua realizzazione ha coinvolto numerosi professori, ricercatori e studenti di vari dipartimenti
dell’ateneo durante il periodo Gennaio 2004 – Luglio 2006. Il lancio del
satellite è avvenuto il 26 Luglio 2006 dalla base russa di Baikonur (KAZ) su un razzo
vettore Dnepr-LV, di derivazione militare. L’orbita prevista sarebbe dovuta essere
una LEO (Low Earth Orbit) con altitudine compresa tra i 600 e gli 800 km per
garantire un deorbitamento autonomo a causa della resistenza aerodinamica dovuta
agli strati alti dell’atmosfera. Il lancio purtroppo è fallito a causa di un problema
idraulico sul razzo vettore.
Il satellite ha la forma di un cubo di lato 13 cm ricoperto su cinque facce da pannelli
solari, fonte primaria di energia per il sistema. Sulla sesta faccia, invece, sono
sistemate le due antenne di comunicazione con terra: una per la banda dei 437 MHz
e l’altra per quella dei 2.4 GHz. All’interno sono presenti sei schede di controllo,
tre telecamere con dierenti distanze focali, sei pacchi batterie per immagazzinare
energia da utilizzare nei periodi di eclissi ed una ruota di inerzia comandata da un
motore che permette il controllo attivo dell’asse di spin del satellite.
Ogni scheda è caratterizzata da una struttura ridondante al ne di garantirne il
funzionamento anche in presenza di guasti.
Le funzioni del satellite sono suddivise nelle varie schede come segue:
PowerSupply
Ha il compito di caricare le batterie utilizzando i pannelli solari mentre la selezione
della batteria da caricare è adata ai due processori di bordo. Si occupa anche del
condizionamento dei segnali analogici provenienti dai sensori di bordo.
Powerswitch
PowerSwitch è la scheda che si occupa di generare le varie alimentazioni per i sottosistemi
del satellite prendendo energia dalle batterie. Su questa scheda sono inoltre
presenti due microcontrollori che si occupano di attivare i processori di bordo e
contare gli eventi di latch-up.
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ProcA e ProcB
Le due schede fungono da processori di bordo e sono sostanzialmente simili nelle
funzioni svolte, ma dierenti nelle soluzioni realizzative. Le principali operazioni che
esse svolgono sono: l’acquisizione dei sensori di bordo, la creazione dei pacchetti di
telemetria, la gestione della carica delle batterie ed il controllo della scheda Payload.
La scheda ProcB si occupa inoltre del controllo del motore elettrico connesso alla
ruota d’inerzia, unica cosa che dierenzia le due schede.
Payload
Compito della scheda è la gestione delle fotograe che vengono scattate utilizzando
le tre telecamere. Inoltre la scheda si occupa della compressione delle immagini in
formato JPEG e della loro trasmissione ai processori di bordo che si occuperanno
poi di inviarle a terra.
TxRx
La scheda TxRx ha il compito di far comunicare il satellite e la stazione di terra.
Sono previsti due canali di comunicazione half-duplex su bande amatoriali: 437 MHz
per la scheda ProcA e 2.4 GHz per la scheda ProcB.
Il lavoro di tesi è iniziato con lo studio dell’implementazione della scheda PowerSupply,
realizzata precedentemente da un altro studente, per poi collaudarla. Terminata
la fase di collaudo di tutte le schede del satellite si è iniziata la fase di integrazione,
che ha richiesto circa un mese di lavoro.
Durante questa fase si è badato a risolvere tutti i problemi di comunicazione fra le
schede e si sono dovute apportare alcune piccole modiche al software dei processori.
Dopo aver completato l’integrazione del satellite si è giunti alla fase nale di collaudo
del sistema, realizzato questa volta utilizzando solamente la comunicazione radio con
la stazione di terra. Per poter giungere a questa fase è stato poi necessario realizzare
il software per una stazione di terra portatile, utilizzata durante le ultime fasi di
collaudo a Torino e durante il collaudo nale svoltosi a Baikonur. Tale stazione di
terra si interfaccia con un PC con ambiente Windows ed ha tutte le funzionalità
della stazione ssa ma con potenza di trasmissione ridotta. Con poche modiche
il programma realizzato può anche essere utilizzato per gestire la stazione di terra
presente sul tetto del Politecnico.
In questa fase sono stati eettuati molti test sul satellite, ultimo dei quali una
prova sulla collina torinese per vericare il puntamento delle antenne, i sistemi di
trasmissione / ricezione e tutte le funzionalità del satellite.
La struttura di questo documento è la seguente:
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Capitolo 1 descrive il progetto globale, la sua struttura ed i vincoli ambientali cui
il satellite è sottoposto;
Capitolo 3 descrive tale scheda e la procedura di collaudo seguita;
Capitolo 4 descrive la scheda processore di bordo B ed il software realizzato;
Capitolo 2 descrive brevemente i protocolli di comunicazione utilizzati tra le varie
schede e con la stazione di terra;
Capitolo 5 descrive la stazione di terra ssa posta sul tetto del Politecnico e la
stazione di terra portatile, di cui è stato realizzato il software;
Appendici riportano gli schemi elettrici delle schede del satellite di cui si è parlato
ed una breve descrizione dell’ambiente spaziale e del modello di cella solare,
utili per una maggiore comprensione delle scelte che sono state fatte nel
progetto.

Data dissemination is the distribution or transmitting of statistical or other data to end users and is
getting more and more importance as the technology is advancing day by day. There are many
ways organizations use to release data to users. Although the most popular dissemination method
today is the non-proprietary open systems using internet protocols but in order to protect our
sovereignty and copyright of the data we will choose to disseminate data using proprietary way so
that we can avoid unauthorized access to our systems.
Idea is to provide a timed and secure access of resource applications of University nanosatellite
ground station setup to end user so that he can make experiments and collect information from any
location through internet. For this I have a secure online reservation system where user must
reserve his timeslot to access the resource application and also I make sure that he cannot have
access before or after the reserved timeslot. Setup includes a complete hardware system to control
University Satellites to make experiments and collect related information.
Although resource sharing through web is not a new concept but it’s not a simple resource sharing
or information sharing I am providing a complete setup to user to perform experiments of his own
choice on University nanosatellites connecting to ground station without making any destructions.

Star Trackers are devices that provide higher accuracy than other attitude sensors
with the added benets of 3-axis attitude determination. Nevertheless, Star Trackers
are frequently heavy, complex and costly systems that can not be adopted by small
satellites such as the Aramis from Politecnico di Torino, which needs high-accuracy
attitude determination to cover the requirements of certain types of payload.
In this thesis, the state of the art in attitude sensing is described, specially that
of Star Trackers. Then, a preliminary design of a low-mass, low-cost, low-power and
coarse accuracy Star Tracker is proposed to satisfy the requirements of the Aramis
spacecraft.
Dierent available algorithms for identifying the presence of single stars on the
imager plane are analyzed, as well as those for pattern recognition necessary to
ultimately measure the spacecraft attitude. One set of such image processing and
pattern recognition algorithms are chosen for use on board Aramis. Subsequently,
they are tested with the experimental use of the 3D open source planetarium Celestia,
while a parallel test of the image processing algorithms is performed on real
star eld imagery to conrm their capabilities with real-world data.
A scheme is proposed to reduce the amount of false results thanks to the use
of attitude approximations coming from other sensors, through the homogeneous
segmentation of the celestial sphere.
Commonly used methods to produce the desired quaternion output are described,
and nally, an assessment of the performance of the tested algorithms is made.

This thesis project focused on the design of a measurement system able to test
attitude control of the academic AraMiS satellite. AraMiS, acronym for Modular
Architecture for Satellites, is a project started in 2006 at the Politecnico di Torino.
The aim of the project is to realize small satellites with a really modular structure.
Modularity allows a significant decrease in the cost of the project itself and thus
provides the university with the opportunity to become interested in space. The cost
of a space mission is, in fact, the main obstacle facing the common interest in space,
by companies and universities. The idea behind this project is to develop dedicated
interconnected and distributed units, built with commercial off-the-shelf components
(COTS) components, in order to increase fault tolerance and allow a decent
performance degradation, while maintaining the costs at acceptable levels. Small
artificial satellites are generally subdivided according to their weight. In particular, it
is about micro-satellite, when the mass of the satellite is between 10 and 100 Kg, it is
nano-satellite, when the mass is between 1 and 10 Kg, it is called pico-satellite, when
the mass is between 0.1-1 Kg. The most efficient way to reduce the cost of a small
satellite project is to reduce project costs as much as possible and not recurring to
manufacturing. These costs, in fact, represent over 90% of the total budget. Their
reduction can only be achieved through the sharing of design between a large number
of space missions. The re-use of projects is the logic behind the AraMiS project, the
development of a modular architecture consisting of a small number of flexible
modules that can be reused in different missions. Reusing the same module over
several times allows you to subdivide project, qualification, and testing costs, and
reduce waiting times before launch.
The most critical part of low-cost spacecraft project is the almost complete lack of
tests, due for several reason, like complexity, time and cost of test equipment. Hence
the idea to realize a low cost multi-purpose measurement system able to model
mainly attitude control of nano-satellite, but also other parameters. To provide a
wide range of tests on nano-satellite has been chosen to design a measurement system
capable to measure position, forces, magnetic field and temperature.

The small size satellites like ARAMIS platforms are low cost solutions for the creation
of several kinds of satellite applications. ARAMIS is a new approach used
in CubeSat. Its faces are PCBs that realize all basic features of the satellite and
that represent also the physical tiles of the CubeSat. ARAMIS structure can be
implemented for different format of CubeSat (1U, 2U, 3U), it is able to bring on
board small telescopes, antennas and other; nevertheless these are characterized to
have a high power consumption, therefore the faces of CubeSat are not large enough
to mount a number of solar panels able to provide enough power to satellite subsystems.
For this reason, a solution with deployable solar panels might be used. In the
launch phase these deployable structures are closed to avoid vibration. After the
expulsion from the P-POD module, the solar panels’ structure is deployed once in
orbit and in this way the total surface of satellite exposed to the sun is increased,
therefore also the power. The deployable solar panels are a solution that can be
adapted to all CubeSat configurations.
The designing of the deploying mechanism, (that allows to open the solar panels’
structure), and the test board where it is placed, is the target of my work. The thesis
deals with the mechanical as well the electrical part of test board focusing on the
compatibility with ARAMIS structure. Trough Altium Design the PCB has been
developed. First of all the circuit has been described through the creation of a blocks
diagram in order to describe the behaviour of the entire system. This operation has
been followed by the physical realization of the board taking into account the right
positioning of components. All elements that form the PCB are chosen considering
the worst case in order to guarantee correct operation. Attention has been paid at
the deploying mechanism used to detach the fixing wires. They keep closed the solar
panels’ structure, during the launch, before passing the atmosphere and entering in
orbit. The UML language has been used to describe main operations and for the
design phase. All the main blocks have the corresponding UML class. All these
classes are referred to Bk1B6711 and 1B111E sections of the ARAMIS project of
the Polytechnic of Turin.

The work presented in this thesis is dedicated to the development of a remote sensor for HUMSAT which is a development of a nano-satellite constellation .
The HUMSAT system architecture is composed of three segments ; space segment , ground segment , and user segment . the user segment formed by the sensors freely deployed and developed by users and by the facilities that users shall design and construct by their own in order to retrieve and send the data.
Data transmission between sensors and the satellite could be Non-bi-directional , bi-directional and full-bi-directional , this thesis describes the non-bi-directional data transmission so a single sensor transmits continuously the frames that it has generated previously until a spacecraft collects them .
A temperature sensor (NTC thermistor ) has been created in order to take part of the user segment , for this goal it has been designed an electronic circuit and its’ PCB , using mentor circuit designer , the circuit is built of a temperature sensor which is connected with a microcontroller ( MSP430F5438 ) and a transceiver ( Si4464 ) , the sensor will relieve the temperature every a defined period and send it to the microcontroller , the microcontroller will recognize if this temperature is in a certain defined range , if the temperature is out of the range the microcontroller will communicate the transceiver to send the data packet .
Some specifics has to be taken into consideration for the data transmission , frequency band is 401-402 MHz , GMSK modulation , EIR is about +27 dBm , for these specifics the Si4464 transceiver has been chosen , the output power of the transceiver was not sufficient , in order to increase the output power of the transceiver a power amplifier is implemented .
Two RF switches is implemented to isolate the transmission and reception chains .
For low power consumption , it is thought to deactivate the transceiver and the PA when it is not necessary to be active , like while the microcontroller is still reading temperature from the sensor , and if the temperature is in the range defined .
In order to configure the microcontroller , the transceiver , and generally reading and transmitting data , C programming codes is written in IAR Embedded workbench .

The small satellites like the ARAMIS platforms [1] are solutions growing low cost and small size, to realize several kind of satellite applications. The ARAMIS is a new approach to the CubeSat designing. It is characterized by the particular designing of the side faces of a CubeSat. These implement PCB boards that realize all the basic features of a satellite platform as reaction wheels, magnetorquer, power management systems and so on. Thus, the PCBs represent also the physical tiles of the lateral faces of the CubeSat. In this way the inner space of the CubeSat will be completely employed all to accommodate a payload. The ARAMIS systems are based on a completely modular approach. All tiles are designed to be easy assembled with interfaces of interconnections realised in standard way. An ARAMIS structure can also be implemented using different format of CubeSat as 1U, 2U, 3U [1], or other more complex formats as the 2x2x2U [1]. They are able to bring on board, payloads like small telescopes or cameras, antennas, remote sensing instruments, small radios telescopes and so on. Many of these payload applications are characterized by a high working power consumption. Just think a remote sensing application that uses ecodoppler techniques to trace the altimetry profile of the earth surface or atmospheric moisture and ionization ones. These instruments usually require high power transmitters and receivers equipment, in order to send and receive the eco signal. So, for these particular high power instruments the only outer tile surfaces of a 1U ARAMIS CubeSat, cannot be large enough to mount a number of solar panels able to provide enough power to the payloads and the satellite subsystems. Furthermore there are situations where the radiation efficiency of the sun is low. It is the case of a space mission much farther from the sun.
So, in all these situations where a large surface of solar arrays needs, they are very useful the arrays of solar panels mounted on deployable mechanicals structures that at launch, for reasons of space, are closed filling a small space. Once in orbit, the structure is deployed increasing the total satellite surface exposed to the sun. In this way, more surface of solar panels can produce more power to supply high power consumption applications.
The main application of the deployable structures if for use in space. Launch vehicles are limited in space and every other kilogram of weight added represents a problem for the launch, mainly in terms of costs. Since a space application is characterised by the not possibility to repair a system, the high cost of a failure leads the space industry to be conservative in the use of its applications or devices, including deployable mechanical structures. In this way, for each for each structure of new concept, the space industry is prevented to the immediate qualification or validation the new systems.
For the market of the nanosatellites nowadays we have been realizing a large amount of deployable structures of solar panels. These structures are deployed in different ways. Generally an elastic mechanical element is used to charge a structure that is suddenly opened when sealing element is released. This is the case of the system developed for example by the Clyde Space company [2].
The designing of a deployable mechanical structure of solar panels compatible for ARAMIS platforms is the target of my work. The thesis discusses in details the mechanical and electrical design, the compatibility of the shape for ARAMIS, its simulations and tests. In addition are explained in details: how the problems of spaces are solved, the electromechanical deployment system and the choice of the employed materials. Are further provided an analysis of opening for a structure of three elements, its thermal analysis and its orbital spin analysis.
The thesis covers also the designing of a board of test for the management of the opening control for a deployable solar panels structure. A specific board is designed. The control circuit of the opening phase is designed on the outer plate element of a reaction wheel tile. In this way, this element will implements on
iv
board also part of the mechanical system of opening of the deployable solar panels structure. The outer plate element is chosen because it is a worst case for the realization of a support tile for the deployable solar panel structure, since it has only one side that can be covered by components and it presents several holes that reduce the space for the circuits. Finally the outer plate board represents a reference model to implement the compatibility interface of the 1B111E, on others different types of ARAMIS tiles. About opening system, particular attention is done to a system of electrical thermal fusers used to detach the fixing wire that maintains the tails of the deployable structure folded during the launch and before the in orbit deployment. The last part of the discussion deals the management software of the outer plate board.
The UML is use to describe the main operation and the design phases. All main blocks are introduced with the corresponded UML class. All these classes are related to the 1B111E and the Bk1B213A1 sections of the ARAMIS project [1] of Polytechnic of Turin.

as, astronomy, atmospheric studies, communications, navigation,
reconnaissance, remote sensing, search and rescue, space exploration, and weather. All
of them provide a great deal of information in the sciences of earth and life.
Usually, a normal sized satellite weighs more than 500kg, it is a few meters long
and can cost more than 1million Euro to build and launch it.[1].
In later years, small satellites (MicroSat, NanoSat, CubeSat and PicoSat) have
been present as the beginning of the space age.
Nowadays, due to advances in microelectronics (especially in microprocessors
and the lower cost on the launching, small LEO (Low Earth Orbit) satellites are an
attractive alternative feasible instead of traditional big GEO (Geostationary Earth Orbit)
satellites.
Small satellites are cheaper to construct and launch, they can do several specific
tasks even better than large satellites, aside from the small size of the satellite.
Furthermore, all onboard components are energy thrifty, and their entire surface is used
to collect energy from the sun through solar panels.
All these advantages provided for small satellites compared with big satellites
can be summarized with the slogan “faster, smaller and cheaper”.
The difference between MicroSat, NanoSat and Pico Sat is basically the size of
each one. CubeSat is a a type of miniaturized satellite for space research with small
volume and mass (less than 1 liter and 1,33 kg). NanoSat is a satellite slightly bigger
than a CubeSat, with a mass between 1 and 10 kg [2].
9
One of the most important parts in a satellite is the Attitude Determination and
Control Subsystem (ADCS). It consists of a set of attitude sensors, actuators and a
microcontroller with control algorithms.
To determine attitude, several sensors can be used: Sun sensor, Horizon scanner,
magnetometer, Star tracker, and Gyroscope. Some of these can be expensive, others
must work together because acting as standalone devices they cannot obtain
adequate/reliable results, and others are too complex or too big for small satellites.
Sun sensor is one of the most common parts in an ADCS. It determines the
satellite’s orientation relatively to the sun by measuring the amount of light that hits it.
It also can be used for instrument pointing, solar array pointing, and safe mode or sun
acquisition systems. There are several types of Sun Sensors, analog or digital, one or
two-dimensional position, all of them, with different specifications and cost. In this
project, a CMOS sensor based is used as Sun sensor.
1.1. Design Goals and Proposed Solutions
The main goal of this thesis is to estimate the attitude of a small satellite using
only image sensors in a low cost satellite.
The main objective is to process the images coming from the image sensor,
which is designed with a low cost Sun sensor that can be used also as a star tracker to
get the attitude of a small satellite (modular architecture, type ARAMIS) without use
any other attitude sensor.
There are two possible ways to solve the problem of the attitude in the satellite
using a sun sensor; the first one is based in the knowledge of the orbit parameters, along
with the measurements from the magnetometer to estimate the rotation angles. This is
the classical solution.
The second approach is presented in this thesis, using only image processing
without knowing the orbit parameters or using any other sensor.
10
Another goal of this thesis is to design and evaluate a real sensor structure for
nano satellites using CMOS image sensor, as a attitude sensor and to evaluate the
possibility to estimate the position of a small satellite using only image sensors.
Such a direct approach can allow for implementations of single sensors to
estimate the attitude, and to reduce cost or space in the satellites.
1.2. Main Contributions
The first period of the PhD program was invested in the study of the literature
and state-of-the-art on the field of Sun sensors for nanoSat, complemented with the
studies of the physical implantation CMOS sensors and design of the Sun sensor for the
satellite. This was an essential part, in order to ensure a proper knowledge base before
starting working towards the aforementioned goals.
The following step was to dedicate the time to design, implementation of the
circuit and board of the Sun sensor, as a part of the modules in Aramis.
Once the schematics circuit was designed, it was changed and redesigned
because the sensor provider (Kodak) in its bankruptcy process cancelled the production
of new sensors; consequently, another sensor from a different provider had to be
selected.
The work with the circuit and the sensor consisted in capture images from the
Sun to calibrate the sensor and to adjust the parameters when neutral filters were used.
The last part of this thesis was done at Space System Laboratory at MIT
(Boston, MA. USA), where the mathematical analysis and captured image simulator
were developed to estimate the attitude of the satellite based in the measurements with
the real Sun sensor (image sensor).

The main scope of this thesis is to prepare the base for the use and the radiation charac-
terization of the new Texas Instruments’ FRAM micro-controllers within the Modular
Architecture for Satellites (AraMIS) developed by the Politecnico di Torino. These kind
of micro-controllers seem to be very appealing for space applications based on Commer-
cial O The Shelf (COTS) components because of their intrinsically radiation hardened
structure and their low power consumption compared with standard FLASH based one.
The idea of using ferroelectric materials to store digital can be dated back to 1952, but
it was practically implemented only starting from the 80s because the needed advanced
technology to develop them wasn’t available before. FRAM based micro-controllers are
instead available on the market since about one year and an half. The ferroelectric RAM
memory, known as FeRAMF or FRAM, is conceptually similar to the DRAM cell, but
there is an important dierence that lies in the dielectric of the storage capacitor: while
DRAM cells use a layer of standard linear material, the dielectric of a FeRAM cell is
made of ferroelectric material, usually lead (Pb) Zirconate Titanate (PZT).
Using a ferroelectric dielectric leads to a dierent behavior of the cell compared
with a DRAM one, leading to many advantages especially for what concern the overall
power consumption in read/write cycles. Furthermore, the material exhibits two stable
polarization conditions and it’s possible to switch between them by means of an electric
eld with opposite polarity. Since the polarization will be kept after the applied eld
is removed, it is possible to link the polarization state to a logic state and so these
materials can be used to build a non volatile memory device. No periodic refresh is so
necessary to keep the information, like in a DRAM memory.
The reading process is destructive: it is not possible to read the content of a cell
without actually clearing it, because of the way the information is stored in the device.
iii
To know which of the possible polarization states the dielectric holds, the only way is to
write a new value to the cell with the bit-line pre-charged but in high impedance state
and depending on the previous polarization, this process will or won’t produce a voltage
pulse out of the bit-line. Read and write cycles require basically the same operations
and can both be completed in times in the order of tens of nanoseconds and without
using high voltage charge pump like in FLASH memories.
The three main design parameters of the electronic systems of small satellites are:
 power consumption;
 physical dimensions;
 radiation environment behavior.
The electric power in the satellite comes from solar panels, which are necessarily of
small dimensions because of the mechanical structure, leading to few Watt of average
power to cover all the needed functions. It is so necessary to make the best use of any
mW of available power. Furthermore launch costs are directly proportional to the mass
of the system, so it is absolutely necessary to reduce as much as possible dimensions
and mass of the electronic system. We said that FeRAM memories are RAM devices,
meaning that read and write procedures do not dier signicantly and random write is
possible without the need of a previous erase of a cell, but they are also non volatile,
we can so for sure state that this leads to save power. In fact in DRAM devices most of
the power is used by the refresh procedure otherwise the stored informations are lost.
Furthermore the refresh process leads to a decreasing in the overall speed performances.
At the moment there are no big FRAM memory available on the market, but in any
case, memory requirements of small satellites are normally compatible with the size of
available FeRAM, except for imaging payloads if local storage of a certain number of
images is mandatory.
Because the FRAM cell stores the state as a PZT lm polarization, an alpha hit
have a very small possibility to cause a change in the polarization. FRAM terrestrial
Soft Error Rate (SER) is not even measurable. This “radiation resistant” characteristic
of FRAM makes it attractive for use in several medical applications and space one.
Arturo Guadalupi iv
Keeping in mind the concepts exposed above, this work is focused on the development
of a payload tile for the AraMIS structure called 1B521 Radiation Characterization
Payload whose aim is to introduce the usage of new FRAM micro-controllers within
the AraMIS nano-satellite structure and characterize them for low cost space applica-
tions in therms of radiations. In particular it is requested to characterize the use of
a FRAM micro-controller (MSP430FR6989) in therms of Total Ionizing Dose (TID),
Single Event Eect (SEE), like Single Event Upset (SEU) and Single Event Latch-up
(SEL), power eciency and reliability in general. No scientic data coming from real
space experiments or terrestrial simulations (using for example particles accelerators)
are available at the moment. It is furthermore requested to show the eciency of the
AraMIS’ developed software hardening library in order to have a direct comparison be-
tween a standard compiled code and an hardened one.
The AraMIS radiation-hardening technique, is based on the use of appropriate C++
classes from the hardened data (Hdata) package developed in house, which can be used
in a common C++ program instead of standard data type. For instance, a short can be
substituted by the so-called TripleShort, which automatically and transparently stores
three copies of the same value and votes or recovers data whenever required. A normal
C++ program can so still be compiled by modifying only the data type denitions.
This makes possible to reuse software algorithms and procedures which have already
been validated and tested without any specic eort apart from redening data types
drastically reducing the development time.
This thesis has to be considered an user guide manual about the developed payload
tile, and a base for future developments on FRAM microcontrollers within the AraMIS
nano-satellite structure. The rst part of the developed work in-fact makes possible
to introduce and start using any kind of FRAM micro-controllers that belongs to the
family MSP430FRxxxx without an heavy eort. All the hardware-dependent choice
that have been made are explained and the software commented in order to be easily
understandable and useful for feature developments.
Arturo Guadalupi v
Here a little overview about the structure of the thesis.
Chapter 1 gives an introduction about the space radiation environments, its interaction
with the electronics and the used shielding techniques.
Chapter 2 gives an overview about the FRAM technology and some concept about
their pro and cons about their use in space applications.
Chapter 3 and Chapter 4 give an overview about the UML approach in the AraMIS
structure and how it is organized.
Chapter 5 explain the design of the developed PCB and what hardware has been chosen
in order to give support to the developed software.
Chapter 6 shows the software structures behind the designed tile, how to it commu-
nicates with the OBC and the type of tests that are executed.
Chapter 7 gives an overview about the tests that has been made to validate the work
and the reached results. What can be done in the feature to improve the what have
been done is also mentioned.