i İSTANBUL TEKNİK ÜNİVERSİTESİ ★ UÇAK ve UZAY BİLİMLERİ FAKÜLTESİ Nanosat ADCS Design and Performance Analysis UNDERGRADUATE THESIS PROJECT Veysel Abdullah TEKİN (110140156) Department of Aerospace Engineering Thesis Advisor : Professor Doctor Alim Rüstem ASLAN February 2021
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i
İSTANBUL TEKNİK ÜNİVERSİTESİ ★ UÇAK ve UZAY BİLİMLERİ FAKÜLTESİ
Nanosat ADCS Design and Performance Analysis
UNDERGRADUATE THESIS PROJECT
Veysel Abdullah TEKİN
(110140156)
Department of Aerospace Engineering
Thesis Advisor : Professor Doctor Alim Rüstem ASLAN
February 2021
ii
İSTANBUL TEKNİK ÜNİVERSİTESİ ★ UÇAK ve UZAY BİLİMLERİ FAKÜLTESİ
Nanosat ADCS Design and Performance Analysis
UNDERGRADUATE THESIS PROJECT
Veysel Abdullah TEKİN (110140156)
Department of Aerospace Engineering
Thesis Advisor : Professor Doctor Alim Rüstem ASLAN
February 2021
iii
Veysel Abdullah TEKİN, student of ITU Faculty of Aeronautics and Astronautics
student ID 110140156, successfully defended the graduation entitled “NANOSAT
ADCS DESIGN AND PERFORMANCE ANALYSIS”, which he/she prepared after
fulfilling the requirements specified in the associated legislations, before the jury whose
signatures are below.
Thesis Advisor : Prof. Dr. Alim Rüstem ASLAN …………………
İstanbul Technical University
Jury Members : Prof. Dr. Cengiz HACIZADE …………………
İstanbul Technical University
Dr. Öğr. Üyesi Cuma YARIM …………………
İstanbul Technical University
Date of Submission : 1 February 2021
Date of Defense : 8 February 2021
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To Dr. Öğr. Kemal Bülent Yüceil
and Halit Ayar
v
FOREWORDS
First of all, I am very happy and proud to have graduated from Technical University. The
meaning of taking lessons from the Faculty of Aeronautics and Astronautic’s
Academicians is that The privilege that was only able to get 80 people each year in
Turkey.
My process of determining the subject of my graduation thesis developed as follows. I
met with aerospace engineering during the Introduction to Aerospace Engineering lecture
opened by Dr. Öğr. Cuma YARIM. Due to the content of the course, I got acquainted
with topics such as orbit mechanics, imaging techniques and satellite technology thus I
started to lay the foundations of what I want to study in the future.
The late Dr. Öğr. Kemal Bülent YÜCEIL had a very special place in me as well as
everyone else in the Faculty. I learned orbital mechanics in depth from Dr. Öğr. Kemal
Bülent YÜCEIL, He was a great scientist. In this way, I started to create the infrastructure
for the location. May Allah raise your soul to the highest heights!
In the 4th year, I took the Attitude control and determination systems course from Prof.
Dr Cengiz HACIZADE. Thanks to the course, I understood the key role and working
principles of Attitude Determination and Control in order to perform the missions of all
spacecraft.
Finally, In Spacecraft Systems Design course, which is considered a undergraduate
course. In this course, Prof. Dr. Alim Rüstem ASLAN made us work with the content
that we could use what we learned in 4 years and feel like a real Aerospace Engineer.
Working with Prof. Dr. Alim Rüstem ASLAN was unique experience as an aerospace
engineer candidate. During this period, I followed his work from the press and felt proud.
Thank you very much for everything.
To sum up, I am grateful for all your efforts and dedication.
February 2021 Veysel Abdullah TEKİN
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vii
TABLE OF CONTENT
Page
FOREWORDS v
TABLE OF CONTENT vii
ABBREVIATIONS ix
LIST OF FIGURES x
SUMMARY xi
ÖZET xiii
1. INTRODUCTION 1
2. ATTITUDE DETERMINATION SYSTEMS 2
2.1 Reference Frames 2
2.1.1 Sun Referenced 2
2.1.2 Central Body Referenced 3
2.1.3 Magnetic Field Referenced 3
2.1.4 Stars and Distanced Planet Referenced 4
2.1.5 Inertial Referenced 4
2.2 Attitude Determination Algorithms 4
2.3 Attitude Determination Systems Sensors 5
2.3.1 Star Tracker 5
2.3.2 Sun Sensor 6
2.3.2.1 Analog Sun Sensor 6
2.3.2.2 Digital Sun Sensor 7
2.3.3 Magnetometer 7
2.3.4 Horizon Sensor 8
2.3.5 Global Positioning System (GPS) 8
2.3.6 Gyroscope 9
2.3.6.1 The gyroscope has two principles 9
3. ATTITUDE CONTROL SYSTEMS 12
3.1 Active Systems 12 3.1.1 Gas Jets/Thruster 12 3.1.2 Reaction Wheels 13 3.1.3 Magnetorquer 13 3.1.4 Ion Thruster 14
ADCS, 2 Ana başlıktan oluşur Yönelim belirleme ve kontrol. Yönelim belirleme uzay
aracının uzaydaki mevcut yönelim bilgisini sağlama sürecidir. Uzay aracının anlık
yönelimi görevin gerçekleşmesi için temel iki girdi den biridir. Kontrol ise mevcut
yönetimden talimatlar doğrultusunda yeni yönelimine geçiş sürecini veya yönelimin
korunmasını ifade eder. Kontrol her zaman aktif şekilde olmayabilir görevin isterleri adcs
çeşidine etki eder. ADCS sürecinde uzay aracı birçok sensör aktivatör ve algoritma
kullanır. Bu ilişki aşağıdaki şemada gösterilmiştir
.
Figure 1 : Closed Loop of ADCS (Markley & Crassidis, 2014, pp. 1–3)
Semaya göre ilk önce uydunun yönelimi için sensör dan alınan veri algoritma ile işlenir
ve algoritmanın sonucuna göre konum belirlenir belirlenen konum doğrultusunda. Emir
oluşturulur veya otonom oluşur. Böylece kontrol sisteminin yazılım aşamasına geçer
burada oluşturulan komut aktivatör aracılığıyla dinamige dönüşür ve closed loop tekrar
başlar. Örnek vermek gerekirse TRYAD cubesat i fırtınalarda ki Gamma işini miktarını
ölçmek ile görevlidir. Görevi gereği tespit edilen fırtınaya doğru yönlenlenmeli ve fırtına
sürecinde bölgeyi hassas şekilde izlemelidir. Bu sebepten yüksek hassasiyetli sun sensör
ve rate gyros kullanır böylece hassas bir şekilde attitude determination yapabilir.
Sözkonusu bölgeyi yüksek hassasiyetle takip edebilmek için geriliği itibariyle gözlem
görevli uzay araçlarında kullanılan reaction wheel ile yonelimini kontrol edebilir. Attitude
determination sistemleri hassasiyet derecesine ve referansa göre değişirken kontrol
xiv
sistemleri aktif ve pasif olmak üzere iki başlıkta incelenir. Öncelikle Atatürk
determination sistemleri şunlardır; star tracker, sun sensor, Earth/Horizon(body
centered), Magnetometer, Global Positioning System GPS and Gyroscope. Bu sistemlerin
her biri farklı amaçlara farklı referans sistemlerine ve farklı isterlere sahiptir. İlerleyen
bölümlerde her sistem ayrıntıları ile işlenecektir. Kontrol sistemlerinde ise aktif ve pasif
olmak üzere iki farklı çeşit vardır. Aktif sistemler sırasıyla, Gas jets, magneto torquer, ion
thruster, reaction wheel and hysteresis rods. Pasif olanlar ise; spin stabilized, gravity
gradient stabilized. (Markley & Crassidis, 2014, pp. 1–3)
1
1. INTRODUCTION
The main purpose of the graduation project is to examine the design and performance
analysis of Attitude Determination and Control Systems in CubeSats. In order to grasp
this process in the best way, the following steps have been followed and made into a
report.
● Literature survey ● Requirements definition ● ADCS elements and design ● ADCS usage for selected missions ● Comparison of designs and effect on operations
After this research and review of literature. 14 CubeSats from different fields and tasks
were selected to embody the work. Each Cubesat's adcs content and tasks were
examined and a trade off table was made. The reason for choosing the ADCS of
CubeSats in 14 different tasks is revealed with a result.
The concept of operation, ADCS and trajectory were determined for the 3 tasks
determined by the analysis performed at the end of the research process. Missions are
given below.
Mission 1: Earth observation, 500 km SSO orbit, mission: take a photo of a 1km*1km
area of a defined location. Determine ADCS requirements, choose suitable ADCS
system, describe how to use the system.
Mission 2: Mars Observation, determine a suitable orbit to place a nanosat with 3
axis ADCS to MAP mars South pole
Mission 3. Sun observation, place a nanosat around earth orbit or sun orbit, to track
sun coronal ejection using a sensor with a field of view of 10 degrees. Selec ADCS
and describe its usage.
2
2. ATTITUDE DETERMINATION SYSTEMS
2.1 Reference Frames
Attitude determination systems are basically classified according to reference systems.
Along with the systems, the sensitivity and the requirements affect the system. To
better understand Attitude-determining systems, we should look at priority reference
systems
Figure 2.1 : References Frames (Wu, 2019)
2.1.1 Sun Referenced
Sun Referenced The determination system, which refers to the sun, determines the
position of the spacecraft according to the angle of incidence of solar rays. In this
system, location determination is highly accurate (potential arc 1 minute). Sun
reference is required to protect the systems from high energy rays and particles coming
from the sun. Sun reference system enables spacecraft to provide efficient energy with
solar cells. In sun referenced determination systems, reference loss may occur due to
the central body. (Wu, 2019)
3
2.1.2 Central Body Referenced
Central body referenced systems require a central mass. The spacecraft / Satellite must
be in or near the central body. The orbiting satellite tracks and defines the earth and
the horizon of the central body optically, therefore the angle of incidence of sunlight
is decisive for the central body (horizon). Its sensitivity is 6 arc minutes. (Hadi, 2015)
Figure 2.1.1 : Central Body Referenced Frame (Hadi, 2015)
2.1.3 Magnetic Field
According to the magnetic field reference, the earth's magnetic field is used in this
system. Sensitivity increases as getting closer to the earth, in LEO, because the earth's
magnetic field decreases from the center to the space. Despite the low energy
consumption in the MFS, the sensitivity is low. (30 arc minutes) In addition, the
systems must be insulated against magnetic effects
4
Figure 2.1.2 : Density of The Earth Magnetic Field (Glatzmaier,
2003)
2.1.4 Stars And Distanced Planet
In order to determine the spacecraft position with respect to the star map reference
system, the position of the stars is identified through optics and their orientation is
determined regarding the assembled map. Additionally, this system is independent of
the orbital and can be used easily in deep space missions. The sensitivity of this system
is very high however the sun rays can mislead the optics of the system. For this reason,
it must be isolated from the sun rays. (Markley & Crassidis, 2014)
2.1.5 Inertial
Source of the inertial reference system is mainly the conservation of angular
momentum. The reference system provides momentum conservation with a gyroscope.
The sensitivity is very high at certain times (when angular momentum is provided) and
it is independent from external reference sources. (Markley, & Crassidis, 2014)
2.2 Attitude Determination Algorithms
Four different algorithms are also commonly used. These algorithms are Geometric
method, algebraic (two vector) method and Wahba method (qmethod). There will be
no detailed discussion of determination algorithms.(Markley, & Crassidis, 2014)
5
2.3 Attitude Determination Systems Sensors
Commonly used attitude determination systems are as follows. They include Star
Tracker, sun sensor, Magnetometer, Horizon sensor, GPS global positioning system
and Gyroscope. (Markley, & Crassidis, 2014)
2.3.1 Star Tracker
Star trackers autonomously make a position estimation in the celestial Frame by
comparing the internal Star catalogs they contain with the Star position detected by
their optical sensors. Their precision is 1 arc second. A star tracker has two modes of
operation: tracking mode and initial attitude acquisition. Tracking mode is mainly the
definition of Star Tracker. It is the process of providing location information in mini
seconds by comparing the images of the spacecraft with its internal catalogs. The initial
attitude acquisition mode, on the other hand, maps the bright star clusters of the
spacecraft used mostly in deep space missions to distant objects and maps the star
reference information in its internal catalog. During this process, tree centroid
calculation is essential for determining location. (Liu, 2011)
Figure 2.3 : Basic Working Principle of Star Tracker (Liu, 2011)
Besides the high sensitivity the sensor provides, it also has a few problems. First of all,
the sensor has a heavy construction, it is an expensive and complex design. The system
consumes high energy. Besides, it must be supplemented with extra sensors and optics
to eliminate the margins of error caused by double stars and multiple systems.
Therefore, a second attitude determination system is often added
6
Figure 2.3.1 : Star Tracker in PicSat (L.E.S.I.A., 2018)
2.3.2 Sun sensor
Sun sensor basically determines the position of the spacecraft by detecting, collecting
and detecting the angle of incidence of the collected rays. Sun sensor is very crucial in
many ways such as angle of incidence of sunlight, energy requirement of the
spacecraft, protection of sensitive components, and protection of sensors using optical
data from sunlight such as Star Tracker. There are two commonly used types of
sensors. Digital and analog sun sensors. (Markley, & Crassidis, 2014)
2.3.2.1 Analog Sun Sensor
It is basically measured by the amount of electric current produced by the sun rays
falling on the photocell. The angle of the sun rays received in the photocell surface is
used to determine the orientation. It has a conical shape to collect much light and it is
more effective in a wide field of views. (Post et al., 2013, p. 6)
Figure 2.3.2 : Basic Analog Sun Sensor Working Principle (Post et al., 2013, p. 6)
7
2.3.2.2 Digital Sun Sensor
It consists of two main parts: command component and measurement component. The
Measurement component gives input to the command component according to the
angle of arrival of the sun rays and it is transformed into input attitude information.
(NASA, 2012)
Figure 2.3.3 : NASA’s Digital Sun Sensor and Its Working Principle Basically
(NASA, 2012)
2.3.3 Magnetometer
It is basically used for the determination of earth's magnetic field for orientation.
Magnetic field lines are from south pole to north pole. Magnetic fields, according to
Faraday principle, create current on the magnetometer and it is used to determine the
direction and density of the current.(Markley, & Crassidis, 2014)
Figure 2.3.4 : Magnetometer Working Principle (ARCBOTICS, 2012)
8
It is a very suitable system for CubeSats in addition to its being low-cost and reliable.
However, it is influential in LEO because of density of Earth's Magnetic Field
2.3.4 Horizon Sensor
The working principle of this sensor is as follows: The spacecraft sees and defines the
horizon of the central body by means of its optics, thus the resulting output creates the
current orientation information. Many of them scan at the infrared wavelength due to
errors or losses in the visible wavelength. Specifically, it works with the same principle
as Star Tracker, but the Reference Frame is the world (for Earth fixed). (Bahar et al.,
2006)
Figure 2.3.5 : Horizon Sensor Principle (Bahar et al., 2006)
2.3.5 Global Positioning System (GPS)
As it is known, GPS is a reliable system that has been used for years to determine
location on Earth. It was first used in 1993 to determine the orientation for satellites.
(Tans vector Sensor). The orientation determination is performed by antennas that
receive signals from different GPS satellites. Position information is obtained by
measurement of the carrier wavelength of the GPS signal falling on the antennas. It is
shown in the figure.(Bahar et al., 2006)
9
Figure 2.3.6 : GPS, ADS Working Principle and Signal Catching with Antenna
(Markley, & Crassidis, 2014)
2.3.6 Gyroscope
Gyroscope has acquired conservation of angular momentum as its basic principle.
Thus, the gyro can measure orientation and maintain this amount. It consists of rotors
intertwined in structure, rotating at high speeds. There are two types of common
gyroscope as rate and laser (Markley, & Crassidis, 2014)
Figure 2.3.7 : Gyroscope Working Principle (Britannica,2020)
2.3.6.1 The gyroscope has two principles;
1-The rotor of the gyroscope tending to keep its rotating axis in space steady. If an
external moment applied to the gyroscope in order to change the conditions of these
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axes, then the gyroscope would resist rotating and in this case frames of suspension
would rotate.
2-The other, second property of the rate gyroscope is called “precession”. Precession
is a special movement of a spinning rotor as a result of external forces acting on it.
Precession refers to a change in the direction of the axis of a rotating object.(Markley,
& Crassidis, 2014)
Figure 2.3.8 : Gyroscope’s Precession Principle Formula
11
Table 2: Attitude Determination Systems Trade-Off table (CubeSatShop, 2009)
Attitude Determination Systems Trade-Off table
Type and Brand
Accuracy Referan
ce Source
Power Reqirements
Orbital Indepen
dence
Mass
Price
Disadvantages
Analog Sun Sensor (NCSS-
SA05)
1 arc minute
Sun 50 mW Yes 5 g 3300 $
1-Because of the Central Body, Input might be lost Digital Sun
Sensor (NFSS-411)
1 arc minute
Sun 150 mW Yes 35 g
12000 $
Star Tracker (STELLA star
tracker)
1 arc second
Stars 250 mW Yes 170
g
>30000
$
must be protect to Sun
Multiple star might be problem
Magnetometer (NSS
Magnetometer)
30 arc minutes
Earth' Magnetic Fields
750 mW No, must
be in LEO
85 g
15000 $
must be toleranced to
magnetic field Work only LEO
Horizon ( OETI)
6 arc minutes
Central Body
(Earth) 50 mW No
100 g
16000 $
must be protect to Sun
Gyroscope (GG1320AN Digital Ring
Laser Gyroscope)
1 arc minute
Inertial 1,6 W Yes 474
g
>20000
$
contain many moving part
need external sensor
determine exact attitude
Global Positioning System GPS
The (NSS GPS Receiver)
6 arc minutes
Central Body
(Earth) 1,5 W No
500 g
- might be connection problem with
GPS satellites
12
3. ATTITUDE CONTROL SYSTEMS
Attitude Control Systems are provided in two forms, active and passive. During the
space mission, it may need to change its orientation to perform the spacecraft mission.
Orientation systems allow the orientation to be changed or maintained actively and
passively. Active systems provide the new position by using energy with actuators in
the orientation determination process or maintain their current conservation. Passive
systems, on the other hand, do not use energy for orientation determination, they either
physically perform the process or orientation is set free.
Active Systems: Gas jets, Magnetorquer, Reaction wheels, Ion and electrical thruster.
Mission:The nanosatellite Chasqui I research project is an effort to secure Peru's
access to space, along with the previous launched satellites, and gives the opportunity
to open new application areas specific to its own geographical and social reality. It is
also from an academic point of view a tool that facilitates collaboration between the
various faculties of the university trains students and teachers with real world
experience in satellite, allowing technological advances in the aerospace industry in
the country. The development of small-scale satellites like Chasqui I gives way to the
various opportunities of access to space with lower costs and development time. For
this reason, various universities, companies and government organizations in the world
show interest in developing nanosatellites that allow to carry out experiments and
scientific missions. The educational benefits of the project can be emphasized in
training camp for future engineers and scientists. (Peru’s National University of
Engineering (UNI), 2014)
ADCS system:magnetometers, sun sensors, GPS and gyroscopes,magnetorquers
Name:SkyCube
Mission:SkyCube had three major mission components: the broadcast of messages
from its radio, the capture of pictures from space via its three cameras, and the
deployment of a large balloon. (Southern Stars Group LLC, 2014)
ADCS system:Magnetometer,magnetorquers
Name:Swayam
Mission:Mission Swayam is the first satellite project of COEP's Satellite Initiative
under the CSAT programme. The team consists of students from freshers to seniors
and spans all the engineering disciplines in the college. The project is in a true sense
an interdisciplinary project. The students in this team are selected after a rigorous
selection avs academic work the team members dedicatedly work on this project all
year round to meet the project deadlines. The team can proudly claim to have published
more than 15 research papers in international conferences for last 7 consecutive
years.(College of Engineering, Pune (CoEP), 2008)
22
ADCS system:magnetometers,MEMS gyroscope,magnets and hysteresis rods
Name:Mars Cube One
Mission:Mars Cube One is the first spacecraft built to the CubeSat form to operate
beyond Earth orbit for a deep space mission. CubeSats are made of small components
that are desirable for multiple reasons, including low cost of construction, quick
development, simple systems, and ease of deployment to low Earth orbit. They have
been used for many research purposes, including: biological endeavors, mapping
missions, etc. CubeSat technology was developed by California Polytechnic State
University and Stanford University, with the purpose of quick and easy projects that
would allow students to make use of the technology. They are often packaged as part
of the payload for a larger mission, making them even more cost effective.(NASA Jet
Propulsion Laboratory, 2019)
ADCS system: cold gas thrusters,reaction control system,star tracker
4.2 ADC Analysis Of Similar Cube Satellites According To Their Duties
When the above CubeSats tasks are reviewed, there are tasks such as earth and Mars
observation, atmospheric analysis, scientific and academic studies, testing of new
technologies and detection of high energy particles. The ADC system is selected
according to the requirements of each task. As an example, the aim of the TRYAD
Cubesat mission is to measure the Gamma ray generation in storms that occur around
the world. As part of the task, CubeSat should regularly monitor the storm and turn its
sensors to the storm center. For this reason, we see that a sun sensor is used for energy
and to protect its optics from the sun. For directional control, reaction wheel which has
a sensitive control capability and used mainly by observation satellites, is used.
Reaction wheel and magnetorquer are also usually present in our system.
Magnetorquer and Magnetometer are included in most CubeSat missions due to their
low energy use and high reliability. The O / OREOS mission is an astrobiology
experiment. It aims to monitor the life cycle of microorganisms under space and stress.
As is seen, the O / OREOS CubeSat does not need to determine the orientation, so it
only carries hysteresis rods. Another specific mission, Mars Cube One, is a Cubesat
made for deep space missions. In this mission, it is planned to test the missions for
Mars. In short, mapping, orbit spiritualities and astrobiological tasks are included. As
mentioned in previous tasks, the reaction wheel is highly essential for observation and
mapping. Star Tracker, which is a suitable system for deep space missions, was used.
23
Considering the spacecraft planned to travel in deep space, Star Tracker is unique for
reference determination and position determination. We see that Mars Cube One uses
cold gas thruster for direction control. The reason for this is that the satellite performs
an orbital maneuver due to its mission. Ion thruster and has jets systems help to change
the satellite's location and orbit as well as the location of the satellite. The ExoCube
example is another useful example in terms of understanding the use of ADC systems.
It does the job of measuring the density of hydrogen, oxygen, helium and nitrogen in
the exosphere during its satellite mission. ExoCube uses gravity-gradient stabilization,
which is a passive system for orientation control, magnetorquer, and a mass
spectrometer and ion sensor for orientation determination. As can be guessed, its orbit
passes through the exosphere, so it does not need an accurate directional control to
catch the gases. As can be understood from the orientation determination system, it
uses an ion sensor to detect the gases that it analyzes. Finally, in the STARS example,
we see that CubeSat was sent for a technology test and the ADC system was chosen
accordingly. In brief, CubeSat consists of two parts and the technology that will allow
these parts to be combined and separated in the orbit is tested, hence it uses a 3 Axis
Magnetometer to determine the orientation. There is no system used for orientation
control other than its own test technology. Thus, as we have experienced from similar
tasks, task requirements determine the correct Attitude determination and control
system.
24
Table 4. : Similar Missions’s ADCS table
Similar Missions’s ADCS table
CubeSat’s Name Missions briefly
ADCS systems
Control Determination
nCube To demonstrate ship traffic. Gravity gradient
stabilization 3-axis Magnetometer
AAU CubeSat To capture pictures into the Earth. B-dot, inertial
sun sensor, 3-axis Magnetometer
STARS To verify the TSR (Connected
Space Robot) system. arm link 3-axis Magnetometer
CubeSail To use radiation pressure exerted
by sunlight for propulsion. blade control sun sensor, 3-axis
Magnetometer
O/OREOS To test how microorganisms
survive and adapt to the stresses of space.
permanent magnets, hysteresis
rods -
TRYAD To measure gamma rays produced by high altitude thunderstorms.
Reaction wheels and magnetic
torquers
Magnetometers, sun sensors, rate gyros
STRaND To test the capabilities of a number
of smartphone components in a space environment.
nano-reaction wheels, 8 pulse plasma thrusters
SSTL’s SGR05 GPS receiver
ExoCube To measure the density of gases in
the Earth's exosphere.
gravity-gradient stabilization,
magnetorquers.
mass spectrometer, an ion sensor
ESTCUBE-2
To investigate deorbiting technology plasma brake, the
propulsion system electric solar wind sail
magnetorquers, gyroscopes
sun sensors, magnetometers
PhoneSat 2.0
to use unmodified consumer-grade off-the-shelf smartphones
Magnetorquers,Reaction wheels
Magnetometer,Gyroscope,Coarse Sun Sensor
Chasqui I To launch Peru's Space and
Satellite program magnetorquers magnetometers, sun
sensors, GPS, gyroscopes
SkyCube To broadcast, take pictures and deploy large balloons.
magnetorquers Magnetometer
Swayam To investigate academics for a group of students.
hysteresis rods magnetometers,MEMS gyroscope
Mars Cube One
To operate beyond Earth orbit for a deep space mission.
cold gas thrusters,reaction
control system star tracker
25
5. MISSIONS : ADCS SELECTION
Concept Of Operation, ADCS Analysis, and ADCS Selection for Missions.
5.1 Earth Observation Mission
Figure 5.1 : Earth Observation’s Concept Of Operation
5.1.1 Mission in brief
In the first mission, CubeSat is asked to take a 1km * 1km photograph of a designated
point on the earth. First, the orbital calculation was made using the given parameters.
According to calculations, the satellite passes vertically above the point to be
photographed every 01,34,36,98 (approximately 1 hour 35 minutes). By the sensor
size and focal length calculation thereafter, the area to be photographed and perceived
in an upright position is shown in the calculations.
26
5.1.2 Camera’s Point of View Calculation
Figure 5.1.1 : Point of View Calculation
(70,5 𝑐𝑐𝑐𝑐)2+ (0625 𝑐𝑐𝑐𝑐)2= 50,105725,
=7,078539 cm 0,6352
𝑥𝑥2= 7,0785392
(500 𝑘𝑘𝑘𝑘)2+ 𝑥𝑥2 ⇒ x = 4,503546099 km
Camera’s Point of View in 90° = 81,12767 𝑘𝑘𝑐𝑐2
5.1.3 Selected Attitude Determination and Control System
Attitude Determination Systems: Digital Sun Sensor, Magnetometer ve Horizon
(Earth)
Attitude Determination Systems: Reaction Wheels and Magnetorquer
Possible cost: 58,850 $ ( cubesatshop.com daki fiyatlar dikkate alınmıştır)
5.1.4 Reason for Choice
The task is an observation mission that requires high precision, as mentioned above.
Hence, it is necessary to determine the area with high accuracy as well as the optics to
be turned to a determined point with the same precision. Primarily, the position of the
sun should be known due to the protection of the optics and energy concerns, and this
is provided by the sun sensor. Since earth observation is in question, the area to be
photographed must be recognized by the satellite and accordingly, the Horizon sensor
is suitable. The Magnetometer, a system that has proven itself many times, has been
used during the mission considering the energy and stand-by conditions. For the
27
control, Reaction wheel, which is used by the largest to small observation satellites, is
preferred. This system is used in large systems such as Kepler and Hubble. With
Reaction Wheel, Magnetorquer is used as it is used in many systems. Thus, an auxiliary
and secondary guidance system is available in energy saving and stand-by conditions.
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5.2 Mapping Mars South Pole
Figure 5.2: Mars South Pole’s Mapping Concept Of Operation
5.2.1 Mission in brief
In the given mission, it is requested to map the planet Mars via the southern hallowed
CubeSat. Orbital selection and operation details are left to the thesis owner. For this
reason, 2 different missions were designed.
According to the first scenario, active orientation determination and control systems
were selected same as in the earth observation mission. In the second method, a choice
was made that could be made at much more affordable prices. Under this scenario,
CubeSat will be able to view the 90,000,000 km ^ 2 south polar circle at a time, within
the calculations shown in the diagrams.
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5.2.2 In Scenario 1: Mapping Mars South Pole
Figure 5.2.1 : In Scenario 1: Mars South Pole’s Mapping Concept Of Operation
5.2.2.1 Mission in brief
the CubeSat passes over the Mars South Pole in every 1 hour and 45 minutes. Thanks
to Active ADC systems, it can photograph 3000*3000 kilometers field from many
angles
5.2.2.2 In Scenario 1, Camera's Point of View Calculation
Figure 5.2.2 : In Scenario 1, Camera's Point of View Calculation (Schenk & Moore,
1999)
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(70,5 𝑐𝑐𝑐𝑐)2+ (0,625 𝑐𝑐𝑐𝑐)2= 50,105725, = 7,078539 cm 0,6352
𝑥𝑥2= 7,0785392
(100 𝑘𝑘𝑘𝑘)2+ 𝑥𝑥2 ⇒ x = 9,00709 km
Camera’s Point of View in 90° = 324,51068 𝑘𝑘𝑐𝑐2
5.2.2.3 Scenario 1, Selected Attitude Control and Determination
Attitude Determination system: Digital Sun Sensor, Horizon sensor for Mars, Star
Tracker.
Attitude control system: Reaction Wheels, thruster/Gas Jets.
Possible cost: 155850 $
5.2.2.4 Reason for Choice:
In our mission, high resolution images should be taken for mapping. Since Marsin does
not have a dense atmosphere, an orbit relatively close to Mars can be chosen, thus
allowing a clear picture. Due to the width of the field, CubeSat's camera must
continually change angle. In this way, it can map with photos taken from many angles.
The system that will provide this orientation capability is Reaction Wheels, which has
been emphasized many times. Since polar orbit is in question, Mars passes through the
south pole every 1 hour and 44 minutes. However, if clearer photos are needed, orbital
transfer may be required. For this reason, Thruster has been added to control systems.
Orbital transfer can be made with this system if requested. In the Attitude
determination section, magnetism based sensors cannot be used since there is no
magnetic field of mars. Since the south pole mapping mission is in question, the Mars-
based base of the Horizon sensor system is used. There is no magnetic based auxiliary
Attitude determination sensor, in addition to the possibility of orbital transfer, the
satellite's position must be determined precisely, so the Star Tracker is assumed
suitable for precise and accurate maneuver. A sun sensor is essential for both energy
supply and protection of the Star Tracker and optical sensors.
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5.2.3 In Scenario 2: Mapping Mars South Pole
Figure 5.3 : In Scenario 2: Mapping Mars South Pole Concept Of Operation (Schenk
& Moore, 1999)
5.2.3.1 Mission in brief
As mentioned in the beginning of the Mars mapping mission, a low budget is targeted
in this scenario. In each period, the optical sensor scans an area of 3000 km * 3000 km
5.2.3.2 In Scenario 2: Camera's Point of View calculation
Figure 5.3.1 : In Scenario 2: Camera's Point of View
Calculation (Schenk & Moore, 1999)
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To find Apogee of Ellipse:
(70,5 𝑐𝑐𝑐𝑐)2+ (0,625 𝑐𝑐𝑐𝑐)2= 50,105725, = 7,078539 cm 0,6352
15002= 7,0785392
(1500 𝑘𝑘𝑘𝑘)2+ 𝑥𝑥2 ⇒ x = 16720,960004 km
5.2.3.3 Scenario 2 Selected Attitude Determination and Control Systems
Attitude determination Systems: Horizon sensor for Mars
Attitude control systems : gravity-gradient stabilization
Possible Cost: 16000 $
5.2.3.4 Reason for Choice
As mentioned in my task description, the aim is a low budget and effective solution.
CubeSat will look towards Mars by taking advantage of the gravity difference with its
gravity-gradient Stabilization system. The mass extending from the body, providing
gravity-gradient stabilization at the exact apogee point, will turn the satellite optics
exactly to the south pole of Mars. Thus, Mars South Pole will be displayed for mapping
in line with the calculations made during the transition from each apogee point. In
addition, the lack of atmosphere will be an advantage to take clear photos even at a
distance.
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5.3 Sun Observation
Figure 5.4 : Sun Observation Concept Of Operation
5.3.1 Mission in brief
CubeSat will track coronal ejections in the sun as defined in the diagram. Effective
solutions were also sought in this task. The most effective solution with the most
affordable costs is of vital importance for an engineer. The sun-synchronous orbit
(SSO), frequently used by weather satellites, is constantly facing the sun, with it
crossing the equator perpendicularly. It is a polar orbit. In this way, CubeSat can stay
facing the sun for 24 hours and throughout its life. Thus, it does not have problems in
terms of energy and can make uninterrupted solar observation.
5.4.2 Selected Attitude Determination and Control Systems
Attitude determination Systems: Sun Sensor and Magnetometer
Attitude control systems: Reaction Wheels and Magnetorquer.
Possible Cost: 42870$
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5.4.3 Reason for Choice
Sun sensor is used to track the position of the sun, Magnetometer used for situations
such as backup, stand by, energy problems, Reaction Wheels used for precise
orientation to the sun's position, Magnetorquer used for auxiliary and backup systems.
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CONCLUSION
The graduation thesis is discussed in 3 main sections. Firstly, in-depth information about ADCS, which was reviewed in the literature, was obtained. The aforementioned information was added to the study together with the visuals in summary. Secondly, 14 different CubeSat tasks were selected and the ADCS systems used by these tasks were examined. Tasks and suitable ADCS systems are shown in tabular form. Thirdly, the operation concept was determined for 3 different tasks and the appropriate adcs were selected.
To summarize, Reaction wheel, which is an active attitude control system, has been chosen as it will actively observe a point in the world. For Attitude Determination, sun Sensor was chosen for the protection of optics and energy requirements. The Horizon sensor is the most suitable for the area to be imaged in the world. In addition, a magnetic based sensor and actuator were used to benefit from the magnetic field of the earth for “Mission 1 : Earth observation, 500 km SSO orbit, mission: take a photo of a 1km*1km area of a defined location. Determine ADCS requirments, choose suitable ADCS system, decsrine how to use the system.” For “Mission 2 : Mars Observation, determine a suitable orbit to place a nanosat with 3 axis ADCS to MAP mars South pole”, Two different scenarios are planned. The scenario is similar to a world observation mission. For this reason, Reaction Wheel, which is an active orientation system, is used. The thruster was added as it is a deep space mission and orbital maneuvers may be required. For Attitude determination, a sun sensor and a Horizon sensor have been added, and since magnetic sensors are not available, the task is guaranteed with Star Tracker. In the second scenario, it is aimed to display the south pole of mars in one go, and it is aimed to provide gravity-gradient stabilization, which is a passive attitude control. Thus CubeSat will always be facing Mars. Horizon sensor is planned for Attitude determination. Finally in Mission 3 : “Sun observation, place a nanosat around earth orbit or sun orbit, to track sun coronal ejection using a sensor with a field of view of 10 degrees. Selec ADCS and describe its usage.” Due to the mission context, CubeSat needs to see as much of the sun as possible. Thanks to the Sso Orbit, it can orbit perpendicular to the equator like meteorology satellites and watch the sun without interruption. Reaction wheel is also suitable for aiming at the sun at any moment. A sun sensor is ideal if the position of the sun is to be noticed at all times. Magnetorquer and magnetometer were chosen to take advantage of the magnetic field again.
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