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

viii

3.2 Passive Systems 15 3.2.1 Spin stabilized 15 3.2.2 Gravity Gradient Stabilization 16

4. SIMILAR MISSIONS 18

4.1 CubeSat List 18

4.2 ADC Analysis Of Similar Cube Satellites According To Their Duties 22

5. MISSIONS : ADCS SELECTIONS 25

5.1 Earth Observation Mission 25

5.1.1 Mission in brief 25

5.1.2 Camera’s Point of View Calculation 26

5.1.3 Selected Attitude Determination and Control System 26

5.1.4 Reason for Choice 26

5.2 Mapping Mars South Pole 28

5.2.1 Mission in brief 28

5.2.2 In Scenario 1: Mapping Mars South Pole 29

5.2.2.1 Mission in brief 29

5.2.2.2 In Scenario 1, Camera's Point of View Calculation 29

5.2.2.3 Scenario 1, Selected Attitude Control and Determination 30

5.2.2.4 Reason for Choice 30

5.2.3 In Scenario 2: Mapping Mars South Pole 31

5.2.3.1 Mission in brief 31

5.2.3.2 In Scenario 2, Camera's Point of View Calculation 31

5.2.3.3 Scenario 2, Selected Attitude Control and Determination 32

5.2.3.4 Reason for Choice 32

5.3 Sun Observation 33

5.3.1 Mission in brief 33

5.3.2 Selected Attitude Determination and Control Systems 33

5.3.3 Reason for Choice 34

CONCLUSION 35

REFERENCE 36

CURRICULUM VITAE 41

ix

ABBREVIATIONS

ADCS : Attitude Determination and Control Systems

ADS : Attitude Determination Systems

ACS : Attitude Control Systems

GPS : Global Positioning System

x

LIST OF FIGURES

Page

Figure 1 : Closed Loop of ADCS xi

Figure 2.1 : References Frames 2

Figure 2.1.1 : Central Body Referenced Frame 3

Figure 2.1.2 : Density of The Earth Magnetic Field 4

Figure 2.3 : Basic Working Principle of Star Tracker 5

Figure 2.3.1 : Star Tracker in PicSat 6

Figure 2.3.2 : Basic Analog Sun Sensor Working Principle 6

Figure 2.3.3 : NASA’s Digital Sun Sensor and Its Working Principle Basically 7

Figure 2.3.4 : Magnetometer Working Principle 7

Figure 2.3.5 : Horizon Sensor Principle 8

Figure 2.3.6 : GPS, ADS Working Principle and Signal Catching with Antenna 9

Figure 2.3.7 : Gyroscope Working Principle 9

Figure 2.3.8 : Gyroscope’s Precession Principle Formula 10

Figure 3.1 : Apollo 11’s Thruster 12

Figure 3.1.1 : Reaction Wheel Devices 13

Figure 3.1.2 : Magnetorquer for CubeSat 14

Figure 3.1.3 : Ion Thruster Schema and Working Principle 15

Figure 3.2 : Spin Stabilization Working Principle 15

Figure 3.2.1 : Gravity-Gradient Stabilization Working Principle 16

Figure 5.1 : Earth Observation’s Concept Of Operation 25

Figure 5.1.1 : Point of View Calculation 26

Figure 5.2 : Mars South Pole’s Mapping Concept Of Operation 28

Figure 5.2.1 : In Scenario 1: Mars South Pole’s Mapping Concept Of Operation 29

Figure 5.2.2 : In Scenario 1, Camera's Point of View Calculation 29

Figure 5.3 : In Scenario 2: Mapping Mars South Pole Concept Of Operation 31

Figure 5.3.1 : In Scenario 2: Camera's Point of View Calculation 31

Figure 5.4 : Sun Observation Concept Of Operation 33

xi

Nanosat ADCS Design and Performance Analysis

SUMMARY

Attitude determination and control system consists of 2 main titles determination and

control. attitude determination is the process of providing the spacecraft's current

orientation information in space. The instantaneous orientation of the spacecraft is one of

the two basic inputs for the realization of the mission. Control refers to the process of

transition from the current management to its new orientation or maintaining orientation

in line with the instructions. The control may not always be active. The requirements of

the task affect the attitude determination and control systems type. Spacecraft use many

sensor activators and algorithms in the Attitude determination and control process. This

relationship is shown in the diagram below.

Figure 1 : Closed Loop of ADCS (Markley & Crassidis, 2014, pp. 1–3)

According to the given scheme, first the data received from the sensor for the orientation

of the satellite is processed with the algorithm and the location is determined according

to the result of the algorithm. The order is created or it occurs autonomously. Thus, it

passes to the software stage of the control system, the command created here turns into

dynamite through the activator and the closed loop starts again. For example, TRYAD

cubesat is in charge of measuring the amount of Gamma rays in storms. It should be

directed towards the storm detected due to its mission and should monitor the region

precisely during the storm. Therefore, it uses high precision sun sensor and reaction

xii

wheel, so it can precisely determine and control attitude. In order to monitor the area in

question with high precision, the observation mission can control its orientation with the

reaction wheel used in spacecraft. Attitude determination systems vary according to the

degree of sensitivity and reference, while control systems are examined under two

headings as active and passive. First of all, attitude determination systems are as follows;

star tracker, sun sensor, Earth / Horizon (body centered), Magnetometer, Global

Positioning System GPS and Gyroscope. Each of these systems has different reference

systems for different purposes and different requirements. Each system details will be

covered in the following sections. There are two different types of control systems, active

and passive. Active systems are Gas jets, magneto torquer, ion thruster, reaction wheel

and hysteresis rods, respectively. Passive ones are; spin stabilized, gravity gradient

stabilized. (Markley & Crassidis, 2014, pp. 1–3)

xiii

Nanosat ADCS Design and Performance Analysis

ÖZET

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

10

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

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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.

Passive Systems: Spin Stabilized, Gravity Gradient Stabilized. (Markley, & Crassidis,

2014)

3.1 Active Systems 3.1.1 Gas Jets/Thruster

This system basically works on the principle of momentum exchange. Attitude control

is provided by the systems similar to the liquid fuel rocket engine on the body of the

spacecraft and high speed gas release from the body. Three-axis attitude provides

control. Thanks to the thruster, very precise and sensitive control is provided. This

system is used in the merging of spacecraft, orbit maneuvers, reparation in orbit, and

missions involving landing where precision is very important. However, since the

system control is provided with high speed gas, the amount of fuel affects the

continuity and quantity of the system. In addition, the correct operation of the thruster

is very crucial for the mission, since there are valves and pipes in the system that carry

fuel and high-pressure gas. Otherwise, the whole task could be endangered. Thruster

was used in Dragon Cargo module, Space Shuttle, Apollo missions and Soyuz module.

(Markley, & Crassidis, 2014)

13

Figure 3.1 : Apollo 11’s Thruster (National Air and Space Museum, Smithsonian

Institution, 2009)

3.1.2 Reaction Wheels

It is a system that uses angular momentum as its basic principle. To illustrate this, we

can give helicopters as an example. In most helicopter systems, when the main rotor

rotates, the tail rotor is activated to prevent the effect of 'turning around itself' and if it

does not balance the rotation effect of the rotor, it starts to rotate with the helicopter

fuselage. Reaction wheel works in the same way. For three-axis rotating, 3 or 4 high

speed rotors are rotated. The amount of rotation speed determines the orientation.

Reaction wheel is a system that can work very precisely and consumes less energy. It

can also be applied to all observation satellites of spacecraft that are larger than

CubeSat. As an example, the Hubble Space Telescope and Kepler telescope use the

reaction wheel to control orientation in deep space observations.(Cortes-Martinez &

Rodriguez-Cortes, 2019, p. 113001)

Figure 3.1.1 : Reaction Wheel Devices (Cortes-Martinez & Rodriguez-Cortes,

2019, p. 113001)

14

3.1.3 Magnetorquer

Magnetorquers generate a magnetic field around the satellite that interacts with the

Earth's own magnetic field, thus creating a torque on the satellite. In this way, the

angular momentum of the satellite can be changed and controlled. The system is

usually used with reaction wheels however it can provide three-axis control in smaller

sizes such as CubeSat. Although the sensitivity is not very high, it consumes a little

energy. Thanks to this feature, it becomes an alternative for orientation in cases of

energy loss that may occur in a satellite or spacecraft. (Markley, & Crassidis, 2014)

Figure 3.1.2 : Magnetorquer for CubeSat (ısıspace, 2013)

3.1.4 Ion Thruster

Ion thrusters are mostly suitable for deep space missions, such as interplanetary

missions, for long-term and highly efficient missions because it can provide propulsion

for very long periods. Besides, as there is no drag force in deep space, the spacecraft

can reach very high speeds even if the amount of thrust is small. Ion thruster is an

advancing technology and it is predicted to be a new generation propulsion system.

Basically, the working principle is based on the acceleration of positive ions in the

magnetic field and the rapid removal of accelerated ions from the nozzle. It is similar

to Gas Thruster with this feature. The process is as follows. Gases are sent to the fuel

tank, the gases in the tank are bombarded with an electron gun. The positive ions that

lose their electrons due the collision, the ions accelerate in the magnetic field, and the

negative and positively charged two plates meet. Due to the potential difference that

occurs here, the plasma, which accelerates rapidly, is thrown from the nozzle and

15

impulse is provided. Thus, the orientation is provided by changing the direction and

geometry of the nozzle.

(Edwards & Gabriel, 1993)

Figure 3.1.3 : Ion Thruster Schema and Working Principle (Räisänen, 2012)

3.2 Passive Systems

3.2.1 Spin stabilized

This system is a passive orientation system, it is generally used for academic purposes.

Basically, it does not give a precise orientation control for the satellite. It rotates around

its axis during the mission process by rotating the satellite while it is put into orbit. For

this reason, CubeSats using Spin Stabilized have limited tasks. It is not suitable for

observation and sensitive experiments.(Markley, & Crassidis, 2014)

16

Figure 3.2 : Spin Stabilization Working Principle

3.2.2 Gravity Gradient Stabilization

Another passive control system is gravity-gradient stabilization. The system directly

uses the law of gravity. A mass is extended (boom-shaped) from the fuselage of the

spacecraft towards the central body. As it is commonly known, the force of gravity

increases and decreases exponentially with the square of the distance. Gravity-gradient

takes full advantage of this law of physics. Since the mass attached to the spacecraft is

closer to the central mass, it is exposed to more force. That is why, the spacecraft

constantly faces the center mass. We can compare this to hammer throwing sport. It is

detailed in the diagram below. This system is not precise and slips may occur.

However, it is a cheap and effective solution.(Wisnievski & Blanke,1999)

Figure 3.2.1 : Gravity-Gradient Stabilization Working Principle

17

Table 3. : Attitude Control Systems Trade-Off Table (CubeSatShop, 2009)

Attitude Control Systems Trade-Off Table

Type and Brand Active or Passive

Accuracy

Reference

Source

Power Reqirements

Orbital

Independen

ce Mass Price Disadvantage

Gas Jet/Thruster ( MicroSpace

MEMS-based Micropropulsion

System) Active Highest

can operate

all Enviroment in Space 2 W yes

300 g

>90000 $

Limited due to the amount of fuel

valve and pipe might

cause problem

Magnetorquer (The ISIS

MagneTorQuer board (iMTQ)) Active Low

Earth's Magnetic field 1,2 W no

196 g 8000 $

Low accura

cy

need magnetic

field

Reaction Wheel (The CubeWheel

Medium) Active High

can operate

all Enviroment in Space 1,5 W yes

140 g 7850 $

contains parts that rotate at high

speeds

Ion Thuster Active Mid

can operate

all Environment in Space

there is no commercial example because of the fact

it is still in improvement Low thrust

Spin Stabilization Passive - - no required -

Gravity-Gradient Stabilization Passive

Lowest

Central Body

(Earth)

no require

d no depending on

the design

It permits one

or two orientations

Low accuracy

18

4. SIMILAR MISSIONS

Names, tasks and payloads of nano satellites (less than 20 kg) in the last 10 years that

orient and control on 3 axes, the adcs they use.

4.1 CubeSat List

Name: nCube: The first Norwegian Student Satellite

Mission: The main mission of the satellite is to demonstrate ship traffic surveillance

from a LEO satellite using the maritime Automatic Identification System (AIS)

recently introduced by the International Maritime Organization (IMO) (Riise et al.,

2003, p. 2)

ADCS system:3-axis Magnetometer, Gravity gradient stabilization

Name: AAU CubeSat

Mission: AAU CubeSat's payload is a CMOS digital camera. It has a resolution of 1.3

megapixels and has a color depth of 24 bit. From the satellite's 900 km altitude, the

camera will capture pictures with a resolution of 120x110 m per pixel. The camera-

chip was provided by a company called Devitech which is in Aalborg. The

Copenhagen Optical Group and the mechanical group developed the lens system. The

structure for the lens system is made out of titanium (Alminde et al., 2003)

ADCS system: magnetometer and a sun sensor

Name: STARS

Mission: The STARS will be normally operated under docking condition, and the

mother and the daughter satellites separated away only when the main mission

performed during roughly 30 seconds.The mission begins based on the command from

the ground station. (Kagawa University & Takamatsu National College of Technology,

2005, p. 2)

ADCS system:magnetic torquer and arm link

Name: CubeSail (UltraSail)

Mission:UltraSail is a proposed type of robotic spacecraft that uses radiation pressure

exerted by sunlight for propulsion. It builds upon the "heliogyro" concept by Richard

19

H. MacNeal, published in 1971, and consists of multiple rotating blades attached to a

central hub (Burton et al., 2021)

ADCS system: blade control (and hence the spacecraft's attitude control)

Name: O/OREOS

Mission: The O/OREOS satellite is NASA's first cubesat to demonstrate the capability

to have two distinct, completely independent science experiments on an autonomous

satellite. One experiment will test how microorganisms survive and adapt to the

stresses of space; the other will monitor the stability of organic molecules in space.

(NASA/ARC (Ames Research Center), 2008, p. 6)

ADCS system: passive magnetic attitude control system

Name: TRYAD

Mission: The Terrestrial Rays Analysis and Detection (TRYAD) project uses 2 small

satellites called CubeSats to measure gamma rays produced by high altitude

thunderstorms. The two CubeSats, TRYAD 1 and TRYAD 2, will use a science

instrument to detect and measure the gamma rays. The science instrument is under

development by the University of Alabama in Huntsville in collaboration with NASA

Goddard Space Flight Center. The two CubeSats carrying the science instrument will

be developed by AUSSP students with mentoring from professors. (the University of

Alabama in Huntsville (UAH) & Auburn University (AU), 2015)

ADCS system:Magnetometers, sun sensors, and rate gyros,Reaction wheels and

magnetic torquers

Name: STRaND

Mission: STRaND started out as a feasibility study and mission requirement exercise

in the Mission Concepts team at SSTL, with the aim of answering the question: ‘what

could SSTL do to leverage the explosion in miniature consumer-level technology that

has occurred in the last 10 years?’. At the same time, SSC were developing an

advanced Cubesat bus, so with input from the advanced space system research at the

Surrey Space Centre, the result of the feasibility study and requirements exercise was

an initial mission concept for a rapid, low cost technology demonstrator and

requirements list, including a list of payload candidates, a component make/buy list,

20

high level concept of operations (CONOPS) and mass, power and budgets. (Bridges

et al., 2011, p. 7)

ADCS system:nano-reaction wheels, nano-magnetorquers, SSTL’s SGR05 GPS

receiver, 8 pulse plasma thrusters (PPTs), and a butane thruster. Except for the SGR05

Name: ExoCube (CP-10)

Mission:ExoCube's primary mission is to measure the density of hydrogen, oxygen,

helium, and nitrogen in the Earth's exosphere. It is characterizing [O], [H], [He], [N2],

[O+], [H+], [He+], [NO+], as well as the total ion density above ground stations,

incoherent scatter radar (ISR) stations, and periodically throughout the entire

orbit.(NATIONAL SCIENCE FOUNDATION (NSF), 2013, p. 44)

ADCS system:Environmental Chamber, gravity-gradient stabilization,

magnetorquers.

Name:PhoneSat 2.0

Mission:PhoneSat is an ongoing NASA project, part of the Small Spacecraft

Technology Program, of building nanosatellites using unmodified consumer-grade

off-the-shelf smartphones and Arduino platform and launching them into Low Earth

Orbit. This project was started in 2009 at NASA Ames Research Center (NASA, 2013)

ADCS system: Magnetometer,Gyroscope,Coarse Sun

Sensor,Magnetorquers,Reaction wheels

Name:ESTCUBE-2

Mission:ESTCube-2 is a technology demonstration mission for deorbiting technology

plasma brake, the interplanetary propulsion system electric solar wind sail (E-sail) and

advanced satellite subsystem solutions.

To test the plasma brake and the electric solar wind sail, commonly called Coulomb

drag propulsion, the satellite will deploy and charge a 300 m long wire which is also

called ‘tether’. Interacting with the ionospheric plasma, the plasma brake will be able

to deorbit the satellite from 700 km altitude to 500 km in half a year. (ESTCUBE,

2020)

ADCS system:sun sensors, magnetometers, magnetorquers and gyroscopes.

21

Name:Chasqui I

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.

28

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.

29

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)

30

(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.

31

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)

32

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.

33

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$

34

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.

35

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.

36

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CURRICULUM VITAE

Name – Surname : Veysel Abdullah TEKİN

Place and Date of Birth (city, dd.mm.yy): ESKİŞEHİR 01.01.1996

E-mail : veyseltekin1683@gmail.com

EDUCATION: B.Sc.: İstanbul Technical University, Faculty of

Aeronautics and Astronatics, Astronautical Engineering