a feasibility study for a satellite vhf data exchange

IN DEGREE PROJECT MECHANICAL ENGINEERING,SECOND CYCLE, 30 CREDITS

, STOCKHOLM SWEDEN 2019

A feasibility study for a satellite VHF Data Exchange System (VDES)

JULIAN GRUJICIC

KTH ROYAL INSTITUTE OF TECHNOLOGYSCHOOL OF ENGINEERING SCIENCES

A feasibility study for asatellite VHF Data ExchangeSystem (VDES)

JULIAN GRUJICIC

Master in Aerospace EngineeringDate: November 13, 2019Supervisor: Christer Fuglesang (KTH), Peter Bergljung (Saab)Examiner: Christer FuglesangSchool of Engineering SciencesHost company: Saab AB, TransponderTechSwedish title: En förstudie för ett satellit väldigt högfrekventdatautbytessystem (VDES)

ii

AbstractTransportation across the globe’s oceans increases every year and is expectedto keep increasing in the following decades. Consequently, there is a needto establish communication over the horizon through the Automatic Identi-fication System (AIS) and the Very High Frequency (VHF) Data ExchangeSystem (VDES), still in development, to track and communicate with vesselsall over the globe regardless of the distance from shore.

In this Master thesis a feasibility study for the development of a systemthat fulfils that need is proposed consisting of a Low Earth Orbit (LEO) con-stellation providing VDES communication continuously all over the globe. Asystem engineering approach has been followed, identifying stakeholders andproducing system requirements setting up a framework for the system. Thekey stakeholders were found to be the customers/users, the satellite provider,the satellite operator, the service provider and the payload provider. Further-more, possible use-cases were presented and a system architecture was definedto outline the system, dividing the system into three segments: the space seg-ment, the ground segment and the launch segment.

In addition, design proposals for a satellite constellation and a typical satel-lite in such a constellation were implemented. The satellite constellation wasproposed to consist of 91 satellites at an orbit altitude of around 550 km inpolar orbits of common inclination, this was regarding a minimum elevationangle of 10 degrees. The satellite is recommended to consist of a 6 U CubeSatusing as payload the existing airborne transponder R5A from Saab Transpon-derTech, it builds on the Software Defined Radio (SDR) technology and is tobe further developed for VDES applications.

Moreover, a link- and a data budget were implemented. Different launchoptions were addressed concluding that launching as secondary payload on aride-share mission or as primary payload on a small satellite launch vehicleare the preferable options. A market analysis has been made providing detailson how many AIS/VDES satellites that have been launched into LEO and bywhich service provider, as well as further details on small/nano satellites ofextra interest to this work. A short risk evaluation was also done, identifyingthe most evident risks with developing, operating and disposing the system.In addition, Saab’s potential role in the development of satellite VDES is dis-cussed.

In conclusion to this work it has been shown that it is possible to build aglobal continuous satellite constellation in LEO utilising as payload an SDR-platform to provide VDES services to vessels at open seas.

iii

SammanfattningTransport globalt till havs ökar varje år och förväntas fortsätta att öka de föl-jande årtiondena. Följaktligen finns ett behov av att etablera över horisontenkommunikation genom det automatiska identifieringssystemet (AIS) och detväldigt högfrekventa datautbytessystemet (VDES), under utveckling, för attspåra och kommunicera med fartyg över hela världen oberoende av avståndetfrån land.

I detta examensarbete har en förstudie utförts för utvecklingen av ett systemsom uppfyller detta behov. Systemet föreslås bestå av en låg jordbana satel-litkonstellation som kontinuerligt tillhandahåller VDES-kommunikation överhela världen. Ett systemtekniskt tillvägagångssätt har följts, intressenter haridentifierats och utifrån dessa har systemkrav tagits fram. De viktigaste intres-senterna befanns vara användare/kunder, satellitleverantören, satellitoperatö-ren, tjänsteleverantören och nyttolastleverantören. Vidare lyftes olika möjligaanvändningsområden för systemet fram och en systemarkitektur framställdesvari systemet delades in i tre segment: rymdsegmentet, marksegmentet ochuppskjutningssegmentet.

Dessutom genomfördes designförslag för en satellitkonstellation samt entypisk satellit i en sådan konstellation. Satellitkonstellationen föreslogs beståav 91 satelliter på en altitud på omkring 550 km i polära banor med gemen-sam inklination, detta var gällande för en minimum elevationsvinkel på 10grader. Satelliten rekommenderades bestå av en 6 U CubeSat med den be-fintliga luftburna transpondern R5A från Saab TransponderTech som nytto-last, vilken bygger på mjukvaruradioteknik och är tänkt att vidareutvecklas förVDES-applikationer.

Vidare, implementerades en länk- och data budget. Olika uppskjutnings-möjligheter undersöktes, varav slutsatsen att uppskjutning som sekundär nyt-tolast på ett delningsuppdrag eller som primär nyttolast medhjälp av ett mindreuppskjutningsfordon anpassat för små satelliter var de föredragna alternativen.Även en marknadsanalys har genomförts, där det redogjorts för hur många AIS/ VDES - satelliter som har uppskjutits i LEO och av vilken tjänsteleverantör,samt ytterligare detaljer om små / nano satelliter av extra intresse för arbe-tet. En kort riskbedömning har också gjorts, där de mest uppenbara riskernamed utveckling, drift och undanröjande av systemet identifierats. Dessutomdiskuteras Saabs möjliga roll i utvecklingen av satellit VDES.

Slutsatsen av detta arbete har visat att det är möjligt att bygga en globalkontinuerlig satellitkonstellation i låg jordbana med en mjukvaruradio somnyttolast som tillhandahåller VDES-tjänster till fartyg på öppna hav.

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AcknowledgementsI would like to take the opportunity to thank my supervisors Christer Fugle-sang at KTH and Peter Bergljung at Saab TransponderTech for their invaluableguidance during months of hard work. Without their insights and experiencethis thesis would not have been possible. Furthermore, thank you to everyoneat Saab TransponderTech for being responsive and encouraging. There wasalways someone around to talk to or ask in times of need.

I would also like to thank my family for their unconditional love and sup-port. Lastly, I would like to thank Orlande Biduax who throughout this workalways was supportive and cheering me on, no matter the circumstances.

Contents

List of Figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiiList of Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiiList of Abbreviations . . . . . . . . . . . . . . . . . . . . . . . . . xxList of Symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . xx

1 Introduction 11.1 Background . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 Intended reader . . . . . . . . . . . . . . . . . . . . . . . . . 31.3 Purpose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.4 Research questions . . . . . . . . . . . . . . . . . . . . . . . 31.5 Need and mission statement . . . . . . . . . . . . . . . . . . 31.6 Delimitations . . . . . . . . . . . . . . . . . . . . . . . . . . 41.7 Chosen methodology . . . . . . . . . . . . . . . . . . . . . . 41.8 Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.9 Literature study . . . . . . . . . . . . . . . . . . . . . . . . . 61.10 Organization of thesis . . . . . . . . . . . . . . . . . . . . . . 7

2 Theory 92.1 AIS and the development of VDES . . . . . . . . . . . . . . . 92.2 The satellite segment of the VDES . . . . . . . . . . . . . . . 10

2.2.1 Down-link . . . . . . . . . . . . . . . . . . . . . . . 122.2.2 Up-link . . . . . . . . . . . . . . . . . . . . . . . . . 13

2.3 Software defined radio . . . . . . . . . . . . . . . . . . . . . 132.4 Orbital geometry . . . . . . . . . . . . . . . . . . . . . . . . 14

2.4.1 Keplerian orbit . . . . . . . . . . . . . . . . . . . . . 142.4.2 Earth coverage . . . . . . . . . . . . . . . . . . . . . 162.4.3 Orbital altitude . . . . . . . . . . . . . . . . . . . . . 192.4.4 Space environment in LEO . . . . . . . . . . . . . . . 202.4.5 Orbital perturbations . . . . . . . . . . . . . . . . . . 202.4.6 Possible orbits . . . . . . . . . . . . . . . . . . . . . 22

v

vi CONTENTS

2.5 Satellite communications . . . . . . . . . . . . . . . . . . . . 232.6 Related work . . . . . . . . . . . . . . . . . . . . . . . . . . 25

2.6.1 Space Norway . . . . . . . . . . . . . . . . . . . . . 252.6.2 ExactEarth . . . . . . . . . . . . . . . . . . . . . . . 26

3 System engineering, stakeholders and use-cases 283.1 A system engineering approach . . . . . . . . . . . . . . . . . 283.2 Stakeholders . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

3.2.1 Stakeholder interest . . . . . . . . . . . . . . . . . . . 293.2.2 Identifying key stakeholders . . . . . . . . . . . . . . 35

3.3 Context diagram . . . . . . . . . . . . . . . . . . . . . . . . . 373.4 Use-cases . . . . . . . . . . . . . . . . . . . . . . . . . . . . 383.5 System limitations . . . . . . . . . . . . . . . . . . . . . . . . 40

4 System design and architecture 434.1 System segments . . . . . . . . . . . . . . . . . . . . . . . . 464.2 Functional architecture . . . . . . . . . . . . . . . . . . . . . 474.3 Space segment . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.3.1 Satellite constellation . . . . . . . . . . . . . . . . . . 474.3.2 Payload . . . . . . . . . . . . . . . . . . . . . . . . . 504.3.3 Satellite platform . . . . . . . . . . . . . . . . . . . . 524.3.4 Choice of satellite provider . . . . . . . . . . . . . . . 58

4.4 Ground Segment . . . . . . . . . . . . . . . . . . . . . . . . 604.4.1 Ground station network . . . . . . . . . . . . . . . . . 604.4.2 Satellite operations . . . . . . . . . . . . . . . . . . . 60

4.5 Launch segment . . . . . . . . . . . . . . . . . . . . . . . . . 614.5.1 Launch vehicle . . . . . . . . . . . . . . . . . . . . . 614.5.2 Separation system . . . . . . . . . . . . . . . . . . . 63

5 Market analysis 645.1 Current AIS constellations . . . . . . . . . . . . . . . . . . . 645.2 AIS CubeSats launched . . . . . . . . . . . . . . . . . . . . . 65

6 Results 676.1 Stakeholder analysis . . . . . . . . . . . . . . . . . . . . . . . 676.2 Technical requirements . . . . . . . . . . . . . . . . . . . . . 67

6.2.1 Stakeholder requirements . . . . . . . . . . . . . . . . 676.2.2 Mission requirements . . . . . . . . . . . . . . . . . . 676.2.3 System requirements . . . . . . . . . . . . . . . . . . 70

6.3 Constellation analysis . . . . . . . . . . . . . . . . . . . . . . 70

CONTENTS vii

6.3.1 Inclination of a sun synchronous orbit . . . . . . . . . 736.4 Design results . . . . . . . . . . . . . . . . . . . . . . . . . . 74

6.4.1 Design budget . . . . . . . . . . . . . . . . . . . . . 746.4.2 Link and data budgets . . . . . . . . . . . . . . . . . 766.4.3 Mission design . . . . . . . . . . . . . . . . . . . . . 80

6.5 Risk assessment . . . . . . . . . . . . . . . . . . . . . . . . . 816.5.1 Identified risks . . . . . . . . . . . . . . . . . . . . . 816.5.2 Analysed risks . . . . . . . . . . . . . . . . . . . . . 82

7 Discussion 847.1 Future work . . . . . . . . . . . . . . . . . . . . . . . . . . . 89

8 Conclusion 90

A System requirements 91

B Link and data budgets 97

C Concept of operations 100

D MATLAB code 109

Bibliography 124

List of Figures

1.1 Showcasing the routes of vessels identified by satellites carry-ing AIS receivers, figure retrieved from [4]. . . . . . . . . . . 2

2.1 An overview of the VDES system is shown here, one can seehow the mentioned shore segment, the ship/vessel segmentand the envisioned satellite segment relate to one another, fig-ures were retrieved from [20]. . . . . . . . . . . . . . . . . . 11

2.2 Showcasing the SAT-VDES, one of three segments of the VDESsystem, figure retrieved from [1]. . . . . . . . . . . . . . . . . 14

2.3 Definition of the orbital elements Ω, ω and i as defined inspherical coordinates with an earth centred reference frame,figure was retrieved from Wakker [18]. . . . . . . . . . . . . . 15

2.4 Definition of parameters needed to calculate the coverage of asatellite orbiting Earth, figure was retrieved from Wertz [13]. . 17

2.5 A comparison of LEO w.r.t MEO, figures were retrieved from[17]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

2.6 Space Norway’s latest satellite NorSat-2, which carries a VDESand AIS payload, in ground tracking (limb pointing) opera-tion, figure was retrieved from [8]. . . . . . . . . . . . . . . . 26

2.7 The Iridium constellation in circular near polar orbit at 780km altitude, it also shows the communication interconnectionsbetween the satellites called Inter Satellite Links (ISLs), figurewas retrieved from [32]. . . . . . . . . . . . . . . . . . . . . 27

3.1 A context diagram of the stakeholders and how they affect andare affected by the system. . . . . . . . . . . . . . . . . . . . 38

viii

LIST OF FIGURES ix

3.2 Potential use-cases when the system will be operational andhow those interact with the user, the satellite operator andthe service provider. Pay attention to how the ground-stationnetwork uses IP to connect with the Mission Control Center(MCC), the data center and possibly directly to the user. . . . . 40

3.3 A WBS of the system. . . . . . . . . . . . . . . . . . . . . . . 41

4.1 The concept characterization process that this work has fol-lowed, where the subject has been defined by the vessel seg-ment of the VDES system as explained in section 2.1, the fig-ure has been retrieved from [12]. . . . . . . . . . . . . . . . . 43

4.2 A simplified architecture of the system where the yellow arrowrepresent the user link at VHF, the blue arrows represent ISLsand the red arrow the feed link. The black arrows representground network (IP) links. . . . . . . . . . . . . . . . . . . . 46

4.3 The development of a typical satellite in a LEO constellationfor VDES from decision at WRC to end of life and de-orbitphase. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.4 Gives an understanding of the need of control in relation to thedistance between satellites, figure was retrieved from [56]. . . 48

4.5 A Walker star constellation of six orbital planes, as viewedwhen looking down on the north pole. One can see the counterrotation between planes 1 and 6 and thus the nodal spacing isdifferent than for the rest of the planes, figure was retrievedfrom [22]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

4.6 Saab’s Airborne transponder R5A, figure was retrieved from[48]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

4.7 The CubeSat specification in the framework of small satelliteclassifications. As can be seen CubeSats are classified basedon mass and volume, figure was retrieved from [15] . . . . . . 54

4.8 Three 6 U Dispensers (from left to right) provided by Plane-tary Systems corporation, ISIS and TNSS, figure was retrievedfrom [73]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

6.1 Swath width as a function of orbit altitude for minimum ele-vation angle 0-10 deg, the calculations are shown in AppendixD. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.2 Orbital lifetime as a function of initial orbit altitude and Cube-Sat form factor. Each satellite has the same mass as its respec-tive form factor, figure was retrieved from [100]. . . . . . . . . 72

6.3 Two different performance indicators, CN0

and Eb

N0, for the satel-

lite communication system plotted against the elevation angle.The results are shown in Tables B.1 - B.2 in Appendix B andcalculations are shown in Appendix D. . . . . . . . . . . . . . 77

6.4 When the antenna gains are not kept constant the carrier tonoise density ratio, C

N0, can be seen increasing at first to an el-

evation of 40 degrees and then to decrease. Results are shownin Tables B.3 - B.4 in Appendix B and calculations are shownin Appendix D. . . . . . . . . . . . . . . . . . . . . . . . . . 78

6.5 The time of satellite contact, Tc, and the data amount,DA, col-lected plotted versus the minimum elevation angle, ϵ, in regardto a single vessel. The calculations can be seen in AppendixD and a Table of the results is given in Table B.5, Appendix B. 79

6.6 Showcases a risk classification matrix over the identified risks. 83

List of Tables

2.1 The classical orbital elements and their definition . . . . . . . 15

3.1 Various stakeholders in the satellite VDES system categorizedas defined by Larson [5]. . . . . . . . . . . . . . . . . . . . . 30

3.2 Trade-off table to identify key stakeholders. . . . . . . . . . . 36

4.1 Emphasizes the design choices that were set and the ones whereoptions were available. . . . . . . . . . . . . . . . . . . . . . 44

4.2 Emphasizes all areas of the WBS, see Figure 3.3, where a de-cision regarding approach had to be evaluated. . . . . . . . . . 45

4.3 Constellation design drivers . . . . . . . . . . . . . . . . . . . 484.4 The physical parameters for the envisioned payload from Saab

for satellite VDES, downscaled to fit the form factor of a 3 UCubeSat. Note: Orbit Average Power (OAP), Peak Power (PP). 52

x

LIST OF TABLES xi

4.5 Potential satellite providers for CubeSats collected from [16]for comparison with the payload requirements from Table 4.4that drives the design of the satellite. . . . . . . . . . . . . . . 59

5.1 Shows service companies providing satellite AIS constellations. 655.2 Shows satellites that have been launched providing AIS/VDES

capabilities. . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

6.1 Stakeholder requirements in the satellite VDES system, whereeach requirement has been defined either as characteristic (Char)or capability (Cap). . . . . . . . . . . . . . . . . . . . . . . . 68

6.2 Mission requirements in the satellite VDES system influencedby what has been done by Øystein et al. [10]. . . . . . . . . . 69

6.3 Showcases how the minimum number of satellitesNtot neededfor global coverage changes with orbital altitude H and whichvelocities vc, periods T and slant ranges D0 are to be expectedwhen keeping the minimum elevation angle at zero, the calcu-lations are shown in Appendix D. . . . . . . . . . . . . . . . . 72

6.4 The number of satellites Ntot needed for continuous globalcoverage and how it corresponds to the number of orbital planesNp and satellites in each plane Ns. The minimum elevationangle is kept at 10 degrees. The slant range decreases with in-creasing elevation angle, see again Appendix D. . . . . . . . . 73

6.5 This table gives an indication of what inclinations are neededto establish a sun synchronous orbit for the altitudes in question. 74

6.6 Table showcases the design budget for a VDES satellite usingISIS 6 U spacecraft bus, pay attention the total mass, volumeand power. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

6.7 Satellite communications parameters for SAT-VDE downlinkand uplink between subject and satellite used in link- and databudget. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

6.8 The satellite antenna gain and the ship antenna in relation tothe elevation angle, as according to ITU-R [117]. . . . . . . . 77

6.9 The preliminary mission design is showcased here. . . . . . . 806.10 Showcases the six most critical development and operational

risks for the VDES system of those that have been identified. . 83

A.1 System interface requirements in the satellite VDES systeminfluenced by what has been done by Øystein et al. [10], aswell as M. Bradbury et al. [8]. . . . . . . . . . . . . . . . . . 92

xii LIST OF TABLES

A.2 System functional requirements in the satellite VDES systeminfluenced by what has been done by Øystein et al. [10], aswell as M. Bradbury et al. [8]. . . . . . . . . . . . . . . . . . 93

A.3 System performance requirements in the satellite VDES influ-enced by what has been done by Øystein et al. [10], as well asM. Bradbury et al. [8]. . . . . . . . . . . . . . . . . . . . . . 94

A.4 System constraint requirements in the satellite VDES systeminfluenced by what has been done by Øystein et al. [10], aswell as M. Bradbury et al. [8]. . . . . . . . . . . . . . . . . . 95

A.5 System environmental requirements in the satellite VDES sys-tem influenced by what has been done by Øystein et al. [10],as well as M. Bradbury et al. [8]. . . . . . . . . . . . . . . . . 96

B.1 The link budget using constant antenna gains for downlinkfrom satellite to subject (vessel). . . . . . . . . . . . . . . . . 97

B.2 The link budget using constant antenna gains for uplink fromsubject (vessel) to satellite. . . . . . . . . . . . . . . . . . . . 98

B.3 The link budget for the satellite downlink to the subject (ves-sel), when the antenna gains of the satellite and ship are chang-ing with the elevation angle as described in [117]. . . . . . . . 98

B.4 The link budget for the satellite uplink to the subject (vessel),when the antenna gains of the satellite and ship are changingwith the elevation angle as described in [117]. . . . . . . . . . 99

B.5 The data budget shows how the time of satellite contact and theamount of data collected from a single vessel are dependent onthe minimum elevation angle. . . . . . . . . . . . . . . . . . . 99

Nomenclature

List of AbbreviationsADCC Attitude Determination and Control Computer

ADS-B Automatic Dependent Surveillance - Broadcast

AIS Automatic Identification System

AMA Adcole Maryland Aerospace

AOCS Attitude and Orbit Control System

ASM Application Specific Message

BCT Blue Canyon Technologies

BER Bit Error Rate

C & DH Command and Data Handling

Cal Poly California Polytechnic State University

CDMA Code Division Multiple Access

ConOps Concept of Operations

COTS Commercial Of The Shelf

CR Constraint Requirements

DSP Digital Signal Processor

xiii

xiv NOMENCLATURE

EMSA European Maritime Safety Agency

EO Equatorial Orbit

EPS Energy Power System

ER Environmental Requirements

ESA European Space Agency

FAL Facilitation of International Maritime Traffic

FFBD Functional Flow Block Diagram

FFI Forsvarets Forskningsinstitutt

FOV Field Of View

FPGA Field Programmable Gate Array

FR Functional Requirements

FW Firmware

GMAT General Mission Analysis Tool

GMDSS Global Maritime Distress and Safety System

GNC Guidance Navigation and Control

GNSS-RO GNSS Radio Occultation

GNSS Global Navigation Satellite System

HKC House Keeping Computer

IALA International Association of marine aids to navigation and Light-house Authorities

Id-Ab Identifiability

ID Identification

NOMENCLATURE xv

IEC International Electrotechnical Commission

IMO International Maritime Organization

IoT Internet of Things

IP Internet Protocol

ISIS Innovative Solutions In Space

ISL Inter Satellite Link

ISS International Space Station

ITU International Telecommunication Union

LEO Low Earth Orbit

LNA Low Noise Amplifier

M2M Machine to Machine

MA Multiple Access

MCS Modulation and Coding Scheme

MEO Medium Earth Orbit

MLI Multi Layer Insulation

MSI Maritime Safety Information

MSS Millennium Space Systems

NASA National Aeronautics and Space Administration

NA NanoAvionics

NCA Norwegian Coastal Administration

NSC Norwegian Space Centre

OAP Orbit Average Power

xvi NOMENCLATURE

PNT Position, Navigation and Timing

POBC Payload On-Board Computer

POCC Payload Operations Control Center

PO Polar Orbit

PP Peak Power

PR Performance Requirements

PSK Phase shift Keying

QAM Quadrature Amplitude Modulation

Re-Ab Replace-ability

Re-Dy Relationship dynamics

REQ Requirement

SAR Search And Rescue

SAS Sky And Space

SAT-AIS Satellite AIS

SAT-VDES Satellite VDES

SDR Software Defined Radio

SEE Single Event Effects

SFL Space Flight Laboratory

SHERPA Shuttle Expendable Rocket for Payload Augmentation

SNR Signal to Noise Ratio

SOCC Satellite Operations Control Center

SOLAS Safety Of Life At Sea

NOMENCLATURE xvii

SSDL Stanford university’s Space systems Development Laboratory

STK Software Tool Kit

STM Sea Traffic Management

SW Software

TBD To be decided

TDMA Time Division Multiple Access

TID Total Ionizing Dose

TNSS Tyvak Nano-Satellite Systems

TRL Technology Readiness Level

TT & C Telemetry, Tracking and Command

UAV Unmanned Aerial Vehicle

UTIAS University of Toronto Institute for Aerospace Studies

VDE-SAT VHF Data Exchange Satellite

VDE-TER VHF Data Exchange Terrestrial

VDES VHF Data Exchange System

VDE VHF Data Exchange

VHF Very High Frequency

VTS Vessel Traffic Service

WBS Work Breakdown Structure

WRC World Radiocommunication Conference

List of Symbols

xviii NOMENCLATURE

α Right ascension

δ Declination

ϵ Minimum elevation angle

η Nadir angle

CN0

Carrier to noise density ratio

Eb

N0Energy per bit to noise power spectral density

Gr

TsSensitivity of receiver

λ Earth central angle

λ0 Maximum earth central angle

µ Standard gravitational parameter

Ω Right ascension of the ascending node

ω Argument of perigee

ϕ Density of the atmosphere

ρ Angular radius of Earth

τ Time of perigee passage

θ True anomaly

e Unit vector from satellite to the sun

f Vector of perturbing force acting on satellite

r Positional vector of satellite

v Velocity vector of satellite

A Effective cross sectional area of satellite

NOMENCLATURE xix

a Semi major axis

Ae Physical aperture area

B Ballistic coefficient

BR Data rate

c Speed of light in vacuum

CD Drag coefficient

CR Power density of the solar radiation

D Distance from satellite to ground target

D0 Distance from satellite to true horizon

DA Data amount

E Specific energy of system

e Eccentricity

EIRP Effective Isotropic radiated Power

f0 Carrier frequency

G Gain of an ideal antenna

g Gravitational force equivalent

Gr Gain of receiving antenna

Gt Gain of transmitting antenna

H Orbital altitude

i Inclination

k Boltzmann’s constant

l Wavelength

xx NOMENCLATURE

Lfs Free space path loss

Lp Polarization loss

M Mass of satellite

Np Number of orbital planes around the globe

Ns Number of satellites in each orbital plane

Ntot Total number of satellites in constellation

Pr Receiver power

Pt Transmitter power

r The radial distance from the centre of Earth to the satellite

RE Mean radius of Earth

REq Radius of Earth at the equator

S Swath angle

Sw Swath width

T Orbital period

Ts System noise temperature

Tc Time of contact with satellite

V Velocity of satellite

vc Circular velocity

Chapter 1

Introduction

1.1 BackgroundIn the globalized world we live in transportation is becoming more and moreimportant. The number of products and passengers being transported acrossthe globe increases every year and is expected to keep increasing the followingdecades [1].

Very high frequency Data Exchange System (VDES), is a radio systemthat is currently being developed to replace the Automatic Identification Sys-tem (AIS). The main driver behind the development of AIS was safety, in thesense that it enabled broadcast communication in which vessels could reporttheir identity, position, speed and course to all other equipped AIS vesselswithin range in order to avoid collisions. Since the introduction of AIS moreapplications of the technology have been introduced, among others aid for nav-igation, Search And Rescue (SAR) and fleet monitoring. Moreover, AIS alsooffered support for introducing Application Specific Message (ASM), whichallow authorities to define custom AIS message types [1].

However, due to globalization the growth has been such that the system isbecoming overloaded endangering the system’s primary mission of safety [1].A number of organizations has therefore started the work on VDES, amongothers the International Association of marine aids to navigation and Light-house Authorities (IALA) and the International Telecommunication Union(ITU) (see Section 3.2 ). VDES rather than being an evolution of AIS is asystem consisting of several subsystems, AIS being one of them. VDES sup-ports Very high frequency Data Exchange (VDE), ASM and AIS incorporatingthree different technologies, but with an increased bandwidth in comparisonto AIS [1].

1

2 CHAPTER 1. INTRODUCTION

Figure 1.1: Showcasing the routes of vessels identified by satellites carryingAIS receivers, figure retrieved from [4].

Nevertheless, According to Lázaro et al. [1] 47.15% of the over 130 000vessels [2] carrying AIS transponders are out of coverage of the coastal VDEStransceivers. Although the system is available globally, it suffers from a ma-jor limitation because of Earth’s curvature, which limits its horizontal rangeto about 70-80 km from shore (depending on antenna height and atmosphericconditions). The implication is that VDES traffic information sent out by shiptransmitters is available only around coastal zones or within reach of ship toship. Utilizing satellites in a constellation carrying VDES payloads overcomesthis problem enabling so called over the horizon capability, making it for ex-ample possible for the ship’s identity and position contained in AIS messagesto be picked up by a satellite and then sent through the constellation to near-est ground station for processing and distribution to users of the system [3].This satellite segment of the VDES system will be called the Satellite VDESsystem and will further be investigated in this feasibility study.

Already today satellite AIS is being used for receive only from vesselsthrough several satellite AIS constellations, where maybe the most prominentare ExactEarth’s, ORBCOMM’s and Spire Global’s, see Table 5.1. Trackingof vessels from such a global satellite constellation can be seen in Figure 1.1.

In contrast, the envisioned satellite VDES system is being developed tohave the capability to not only receive but to transmit data to vessels as well.

As Lázaro et al. [1] further explain, a common goal of the terrestrial seg-ment (the combined shore and vessel segment) and the satellite segment of theVDES system is to keep the development costs low and to target a simple re-placement of all AIS transceivers with a single ”VDES box”. This correspondswell with the goals of Saab AB and is in accordance with this study.

CHAPTER 1. INTRODUCTION 3

1.2 Intended readerThe intended readers are Saab employees and fellow students that will furtherdevelop on the idea of Saab going back to Space.

1.3 PurposeThe aim of this work was to investigate if Saab has the prerequisites to incor-porate satellite VDES based on the Software Defined Radio (SDR) - platformdeveloped at Saab.

1.4 Research questions• What requirements are put on the payload and the satellite respectively

for the satellite VDES system?

• Is it possible to use the existing SDR-platform developed at Saab aspayload for the purpose?

• How can global coverage be achieved by incorporating a satellite con-stellation called satellite VDES into the terrestrial VDES system?

1.5 Need and mission statementIt has been identified that there is a need of establishing global coverage, alsocalled over the horizon capability, for tracking and communicating of/withvessels to increase safety at open sea.

This report is to evaluate whether a satellite constellation with an SDRplatform developed at Saab for VDES can fulfil that need.


1.6 DelimitationsThis study is not going to investigate the whole VDES system, but rather focuson the satellite segment of the VDES system, refereed to as the satellite VDESsystem.

This report is not going to investigate all the specific subsystems for thesatellite nor for the VDES-payload in detail, but rather give an overview of thewhole satellite VDES system.

It is not aimed at investigating the possibility of navigation, also calledranging from VDES payload, instead this is left to further studies.

1.7 Chosen methodologyThe methodology was chosen after intensive discussions regarding how to sys-tematically approach the subject. In order to find an answer to the researchquestions a system engineering approach described by Larson [5] was selected.The work has been done as follows:

1. Literature study regarding satellite VDES.

2. Introduction to the background of the subject.

3. Delimitations and assumptions.

4. Investigate which stakeholders exist. Which company can produce thesatellite? Which company can offer launch opportunities? Who is thecustomer? What competitors are there?

5. Investigate possible use-case scenarios. How does the system work todeliver data to its users?

6. Examine signal strength and coverage for the orbital altitude that is de-cided to be used.

7. Examine space environmental aspects such as radiation, air resistance,solar radiation pressure and earth oblateness effect for chosen orbitalaltitude.

8. Develop a requirement specification for the satellite and the SDR plat-form for VDES purposes.


9. Define system architecture. How does the system behave functionallyand how may the system look physically?

10. Develop design budgets for mass, size and power.

11. Develop link and data budgets.

12. Examine the type of constellation that can be used, how many satellitesare required to achieve desired coverage and communication speed?

13. Develop a market analysis to examine the business incentives for Saabto make an entrance into the field.

14. Assess technical risks with the development, operation and disposal ofthe system.

1.8 MaterialThis work has as far as possible used research articles published in peer-reviewedjournals and well established textbooks to reference and build its argumentsupon.

The material is mainly from technical reports and journals retrieved from:

• Google Scholar

• KTH Library (Primo)

• Google books

• SCOPUS

These sources were used since they are easily accessible and can be saved ina larger library for future use.Key search words have been:- VDES- VDE-SAT- AIS-SAT- Satellite Automatic Identification System (Sat-AIS)- Low Earth Orbit (LEO) Constellation- Continuous global satellite constellation- LEO satellite constellation design- Inter Satellite Link (ISL)


- Software Defined Radio (SDR)- CubeSat- Small/nano satellite- (Small/nano satellite)/CubeSat design- Link budget for small satellites- Small/nano satellites trends/forecastFurthermore, citations from already found papers have been very useful inachieving more depth in the analysis.

Regarding the methodology of system engineering, it has mainly been re-trieved from literature described in Section 3.1.

Lastly, newspaper articles, satellite databases such as e-portal, Kulu [6] and[7] as well as data sheets have been used to provide data regarding the currentmarket situation for satellite AIS/VDES and to construct design budgets. Suchsources have many times been double checked to clear doubts of misinforming,nevertheless one cannot evidentially assure that all data is correct.

1.9 Literature studyTo begin with, Space Norway and their ongoing research of tracking vessels bysatellite has been found interesting, especially since they launched the first op-erational satellite carrying a VDES transmitter as payload in 2017. This madeit the first company to test the VDE down-link from their satellite NorSat-2 tovessels carrying VDES transponders [8].

Another company of interest has been ExactEarth, which has one of fewoperational SAT-AIS constellations. It has worked within the field for over 12years, establishing its own constellation ExactView that co-operates with itsSDR payloads on the Iridium constellation to bring global coverage to its users[9].

Lázaro et al. [1] have been an inspiration to why the need for a satelliteVDES system exists and to many of the identified use-cases. They have alsoprovided insight to system constraints and to future work that is needed to beaddressed when constructing this type of system.

Moreover, an interesting study that was of importance to this work was”ESA Mullighetsstudie for et eurpeisk satellitbaserat AIS system” by Øysteinet al. [10], in which a feasibility study for the development of an Europeansatellite AIS system was performed. They faced many similar problems tothis study. In their work they developed system requirements, see pages 78-85 in [10], to which this study and in particular the system requirements, seeAppendix A, have taken influence. What’s more, regarding the system require-


ments, NASA’s System Engineering handbook [11] helped to categorize thesystem requirements and also functioned as an inspiration and framework todevelop the concept of operations featured in the Appendix C.

For most of the design part of this study Wertz and Larson [12] have beenreferenced to motivate design choices of the mission as a whole as well asthe satellite, whereas Wertz [13] has been referenced when conducting thecoverage analysis for the satellite constellation design. In addition, J.Ippolito Jr[14] has been referenced, apart from already mentioned [12], when conductingthe link- and data budget.

Poghosyan and Golkar [15] have analysed current capabilities for Cube-Sats, being a source of information from which the satellite design has beenrelated to. Another source was the NASA’s report ”State of the Art” [16]contributing with insights on CubeSat capabilities and evolved technology.This report also provided perception of current satellite providers, deploymentproviders, launch providers and ground station providers. In addition, datasheets over subsystems of the satellite, retrieved from nano satellite providers’websites, helped to construct the design budgets.

Due to the work of Reid et al. [17] the author has limited the work to onlyconsidering LEO as the better approach for a CubeSat satellite constellationas is explained early on in section 2.4.3.

Lastly, Wakker [18] has been referenced when developing on orbital ge-ometry and orbital perturbations.

1.10 Organization of thesisIn ”Introduction”, see Chapter 1, a short explanation of the background andaim of the report is given.

Moreover, in Chapter 2 ”Theory”, theory regarding the development of thesystem has been brought forward, giving an introduction to AIS and VDES,orbital geometry as well as theory regarding Earth coverage, satellite commu-nications and related work.

Then follows the Method organized into three chapters: First Chapter 3,called ”System engineering, stakeholders and use-cases”, where the systemengineering approach is introduced, stakeholders and use-cases of the satelliteVDES system are identified and lastly system limitations are presented. Inthe following Chapter 4 ”System design and architecture” the architecture anddesign of the system is made, outlining the system and its interface with theuser segment (Terrestrial VDES), then the constellation design requirementsare explained and finally the design choices of each of the system segments


(Space, Ground and Launch) are given. The last part of the Method Chapter 5”Market analysis” provides details regarding how many AIS/VDES satelliteshave been launched and by whom as well as further details on nano satellitesof extra interest to the work.

Furthermore, follows Chapter 6 ”Results”, where the stakeholder, missionand system requirements have been listed, The design budget, link budget anddata budget have been presented, the constellation analysis and a risk assess-ment have been made.

In ”Discussion”, see Chapter 7, the results are analysed and discussed,future studies that are needed for the development of the system are presentedand the author takes the freedom to be a bit visionary for what to come for thefuture of the satellite AIS/VDES industry. Finally, in ”Conclusion”, Chapter8, the Research questions, see section 1.4, are answered as concise as possible.

Chapter 2

Theory

2.1 AIS and the development of VDESThe VDES system is building on the capabilities of AIS addressing the needfor more data capacity. It is seen as an effective and efficient use of radiospectrum by providing higher data rates than AIS and optimizing the protocolfor data communication [19].

Below the outline of the VDES system is described. The VDES system iscomprised of three main segments, the vessel segment which defines all theactive vessels at sea, the shore segment which is provided by the base stationnetwork across the globe and the envisioned satellite segment which is outlinedin this report, working together to keep track of a vessel’s position, velocityand course to avoid collisions [20]. Note that the combination of the shore- andvessel segment previously (see section 1.1) has been outlined as the terrestrialVDES segment.

As earlier mentioned (see section 1.1) the communication architectureof the system is composed of three subsystems called AIS, ASM and VDE.Where the highest priority is given to AIS for position reporting and safetyrelated information and thereafter follows ASM and last comes VDE with theleast priority of the subsystems as stated in M.2092-0 [20].

The major differences between AIS and what the two new VDES sub-systems (ASM and VDE) can provide are identified by IALA in [19] as thefollowing:

1. The Modulation and Coding Scheme (MCS).It is adaptive, choosing the coding rate and modulation according to thecurrent SNR (signal to noise ratio) and identifies the modulation moreefficiently than for AIS [1].

9

10 CHAPTER 2. THEORY

2. The radio frequencies that are used.

3. The increased data bandwidth. Providing higher data rates

4. The methods used by the link layer.

The system operates on the maritime frequency band, where each channelhas a specific frequency. To shortly give an overview of the space segment,the channels are as follows:

• AIS 1 (channel 2087) and AIS 2 (channel 2088) are terrestrial AIS chan-nels that are also used as uplinks for receiving AIS messages by satellite.

• Long Range AIS using channel 75 and channel 76 are specified channelsto be used as uplinks for receiving AIS messages by satellite.

• SAT Uplink 1 (channel 2027) and SAT Uplink 2 (channel 2028) are usedfor receiving ASM by satellite.

• SAT Uplink 3 (channels 1024, 1084, 1025, 1085, 1026 and 1086) areused for ship-to-satellite VDE uplinks.

• SAT Downlink (channels 2024, 2084, 2025, 2085, 2026 and 2086) areused for satellite-to-ship VDE downlinks.

This proposed channel configuration has been described as outlined in Inter-national Telecommunication Union (ITU)’s report on VDES M.2092-0 [20].It is expected to be accepted in November at the World RadiocommunicationConference (WRC) and their allocation can be seen in Figure 2.1b.

Moreover, to clarify in regard to the satellite segment, it is only the SAT-VDE channels that are for satellite downlink to the vessel segment, whereasASM and AIS are for uplink only, as can be seen in sub figure 2.1a.

2.2 The satellite segment of the VDESThis section will explain the concept of the satellite segment for the VDESsystem in more depth. An operative satellite in such a segment, as envisionedby Lázaro et al. [1], can be seen in Figure 2.2.

Each satellite should use a Bulletin Board system and Announcement chan-nels to communicate to vessels in its coverage area for both downlink and up-link [20].

CHAPTER 2. THEORY 11

(a) An overview of VDES.

(b) An overview of the VDES’s channels and frequencies.

Figure 2.1: An overview of the VDES system is shown here, one can seehow the mentioned shore segment, the ship/vessel segment and the envisionedsatellite segment relate to one another, figures were retrieved from [20].


Furthermore, as explained in M.2092-0 [20] the link budget of the down-link from satellite to vessel should take into consideration path losses, propa-gation losses, receiver sensitivity of merit and local interference levels. Nev-ertheless, also the elevation angle in regard to the vessel must be considered,where most transmissions will apply to a vessel at low elevation angles caus-ing a further satellite range, which in return sets requirements for the antennagain.

In order to give an example of this range difference, consider a satelliteat an orbital height of 600 km, depending on the elevation w.r.t the vessel therange will differ from the minimum of 600 km to the maximum of 2830 km,indeed high satellite range variations causing a path delay variation from 2 ms- 10 ms [20].

Moreover, a satellite at that altitude will move at about 8 km/s causing amaximum Doppler shift of about ±4 kHz at VHF according to M.2092-0 [20].This will occur when the relative velocity between the vessel and the satelliteis at its maximum, which happens at low elevation angles.

2.2.1 Down-linkAs earlier stated and shown in Figure 2.1, only the VDE element of VDES(VDE-SAT) is used for downlink from the satellite to the vessel segment.Here follows the services that VDE-SAT downlink should support accordingto M.2092-0 [20].

• Downlink multicast multi-packet data transfer (satellite-ship).

• Shore originated unicast multi-packet data transfer via satellite (shore-satellite-ship).

The network layer that makes this possible, as defined as well in M.2092-0[20]:

• Bulletin Board transmissions (network configuration).

• Multicast (satellite-ship), which gives information to mariners regardingice maps and routes, weather info, etcetera. This information is directedto all ships within the Field Of View (FOV) of the satellite.

• Unicast (Shore to ship transfer, up to 100 kb), this corresponds well withships that have signalized through uplink that they need information,


but are out of reach of the shore segment(shore-satellite-ship). Thisinformation is directed to a specific vessel(s).

2.2.2 Up-linkIt is envisioned for satellite up-link that all three subsystems (AIS,ASM andVDE) are supported, using the following types of functionality according toM.2092-0 [20]. As for the VDE-SAT uplink it can be divided into: Two-waycommunications and Transmit only.

Here Two-way communication (for vessels out of range of the shore seg-ment) corresponds respectively to information that the shore would want fromvessels (response from vessel) and information that vessels would want fromthe shore (request from vessel). This is information is uplinked from the ves-sel and travel through the satellite segment to reach the sought receiver in theshore segment.

Transmit only would be either event driven or periodic. The time slot andfrequency band for this service should be assigned by the bulletin board andannouncement signalling channels. Examples of these kind of transmissionswould be a mayday call, having obviously the highest priority of all messages.

To describe the network layer it will based on:

• Ship originated single packet data transfer.

• Ship originated multi-packet data transfer.

As defined in M.2092-0 [20].

2.3 Software defined radioWhen it comes to communication technologies Software Defined Radios (SDRs)are becoming more popular, replacing conventional radios due to their in-crease in flexibility, data throughput and scientific output [21].

An SDR is basically defined by the signal processing and their algorithmsfor a specific service such as AIS or VDES. Most SDRs are implemented indigital signal processors (DSP) or Field programmable gate arrays (FPGAs)[21].

The main advantage according to Budroweit [21] is the possibility to up-date the baseband processing system. An important aspect when it comes tosatellite operations, since it opens up the possibility to improve performance asthe satellite is in-orbit. What more is, it reduces cost, size and mass by havingmost of the data processing being done in one single chip.


Figure 2.2: Showcasing the SAT-VDES, one of three segments of the VDESsystem, figure retrieved from [1].

2.4 Orbital geometry

2.4.1 Keplerian orbitInitially, consider a Keplerian orbit which follows from Kepler’s three laws. AKeplerian orbit is a conic section, which is defined as the geometric collectionof all points P for which the ratio to a fixed point F and the distance to a fixedline l is constant [18], restricted by the two-body problem concerning the twobodies i and k, where k is the more massive body [18]. For example the moonorbiting Earth or a satellite orbiting Earth, neglecting all other perturbing bod-ies. The position of a satellite (body i) with respect to Earth (body k) can beexplained by using only six independent parameters called the classical orbitalelements [18], see Table 2.1 and Figure 2.3.

Looking at Figure 2.3, one can in addition to the classical orbital elements,also distinguish the true anomaly θ which describes the angle from the perigeeto the point of the satellite in its orbit, r the distance from the centre of Earthto the satellite and lastly there are the astronomical coordinates declination δ


Table 2.1: The classical orbital elements and their definition

Orbital elements Definition Descriptiona Semi major axis The size of the orbit.e Eccentricity The shape of the orbit.τ Time of perigee passage The time passed since last perigee passage.i Inclination The angle between the orbital plane and the equatorial plane.ω Argument of perigee the angle from the node of lines to the perigee.Ω Right ascension of the ascending node The angle from the vernal equinox to the node of lines.

Figure 2.3: Definition of the orbital elements Ω, ω and i as defined in sphericalcoordinates with an earth centred reference frame, figure was retrieved fromWakker [18].

which is the angle between the equatorial plane (reference plane) to the pointof the satellite in its orbital plane and α the right ascension defining the anglefrom the Vernal equinox eastward along the equator to where the orbit of thesatellite crosses the equatorial plane from south to north.

In order to find the velocity and the orbital period of the satellite in itsorbit one may begin with defining the Vis Viva - equation commonly knowwithin the field of Astrodynamics [18]. It describes the specific energy E ofan orbit. In this case, the energy is constant and negative, considering a stableorbit where both bodies stay in the system. Furthermore, it can be shown thatthe specific energy is equal to −µ

2a, see Equation 2.1, for a elliptical or circular

Keplerian orbit [18].

E =v2

2+

µ

r=

−µ

2a, (2.1)


where µ is the standard gravitational parameter and r is the radial distancebetween the centre of each respective body. In the case of a satellite in orbitaround Earth, µ is the standard parameter of Earth and r the radial distancefrom the centre of Earth to the satellite in its orbit.

Assuming a circular orbit a = r makes it possible to determine the veloc-ity of the satellite in its orbit also referred to as the circular velocity vc, seeEquation 2.2.

vc =

√µ

r(2.2)

, where also the period T of the orbit can be determined by Equation 2.3

T =2π r

vc= 2π

√r3

µ. (2.3)

As indicated by Equation 2.2, the velocity will increase with lower orbits.Moreover, to calculate the distance to the satellite from the centre of Earth

the trajectory equation can be used, see Equation 2.4

r =a(1− e2)

1 + e cos θ, (2.4)

here θ is represented by the true anomaly which is the angle measured fromthe perigee to the current position of the satellite as seen in Figure 2.3 [18].

2.4.2 Earth coverageIn order to decide the number of satellites needed for achieving full coverageone can make an estimation by considering the setting showcased in Figure2.4, as depicted by Wertz [13] for one satellite. A few assumptions have beenmade to simplify the analysis:

• Circular orbits are considered.

• Earth has been approximated as spherical, disregarding Earth’s oblate-ness.

• Earth has been assumed to not rotate, if full coverage is reached there isno difference to a rotating Earth [13].

Having taken these assumptions into consideration, earth coverage is de-pendent on the following variables [13]:

• Number of satellites


• Altitude

• Inclination

• Revisit time

• Swath width

Looking at figure 2.4, one can distinguish the different angles and distancesthat are of importance when describing coverage from a satellite: ϵ representsthe minimum elevation angle, n the nadir angle, λ0 the maximum earth centralangle, ρ the angular radius of the Earth, H the orbital altitude, RE the meanradius of Earth, D0 the distance from the satellite to the true horizon, λ theEarth central angle and D the distance to the target as seen from the satellite.

Figure 2.4: Definition of parameters needed to calculate the coverage of asatellite orbiting Earth, figure was retrieved from Wertz [13].

A number of equations have been helpful to decide how many orbitalplanes are needed to cover Earth, this is dependent on the swath angle S whichis an angle that defines how many degrees of Earth is covered by each respec-tive satellite [13]. The swath angle can be decided by the maximum earthcentral angle as in the following equation 2.5

S = 2 · λ0, (2.5)

which in turn is dependent on the chosen altitude as can be seen in equation2.6

λ0 = arccosRE

RE +H. (2.6)


The diameter of the circular ground track can be calculated by multiplyingthe swath angle with the radius of Earth at the equator REq [22]. This is calledthe Swath width Sw, and is shown by equation 2.7

Sw = S ·REq. (2.7)

The number of planes needed, Np, for global coverage can be decided accord-ing to Akyildiz, Jornet, and Shuai [23] by taking into consideration that eachsatellite can cover S degrees of the equatorial region and that all satellites havea common inclination of polar or near polar orbit, as given by equation 2.8,

Np =360

2S. (2.8)

And in a similar manner one can decide the number of satellites in each orbitalplane Ns, by equation 2.9,

Ns =360

S, (2.9)

whereas the total number of satellites Ntot in the constellation can be decidedfrom equation 2.10,

Ntot = Np ·Ns (2.10)

[23]. However, if the minimum elevation angle ϵ is known and not given bythe true horizon ϵ = 0, instead it is constrained by for example the satellite’santenna gain or the target antenna, one has to decide the other parameters fromFigure 2.4 a bit differently [13].

To begin with one can define ρ the angular radius of the Earth as follows

ρ = arcsinRE

RE +H. (2.11)

Using equation 2.11 and considering the known minimum elevation angleone can determine the nadir angle at that specific elevation by

η = arcsin (cos ϵ · sin ρ), (2.12)

and finally by using the result from equation 2.12 one can determine the earthcentre angle at the chosen minimum elevation angle as

λ =π

2− η − ϵ. (2.13)

with the result of 2.13 one can again using equations 2.5-2.10 define the num-ber of satellites needed in the constellation.


Lastly, the distances given from the satellite to the target and the true hori-zon respectively, called slant ranges, are given by equations 2.14-2.15

D = RE(sinλsin η

), (2.14)

D0 =RE

tan ρ. (2.15)

2.4.3 Orbital altitudeWhen deciding upon orbital altitude two options were considered, either LowEarth Orbit (LEO) or Medium Earth Orbit (MEO). As Reid et al. [17] empha-sise there are several key notes for why LEO is preferable over MEO consider-ing Position, Navigation and Timing (PNT) use-cases. The difference betweenLEO and MEO can be seen in Figure 2.5.

(a) A comparison between LEO andMEO, footprint to scale.

(b) Data for typical constellations inMEO (GPS) and LEO (Iridium).

Figure 2.5: A comparison of LEO w.r.t MEO, figures were retrieved from [17].

• Satellites travel faster around earth -> more multipath rejection.

• Lower altitude -> Less path loss and delivering signals up 1000 times(30 dB) stronger than MEO.

• Less path loss and stronger signals -> more resilient to jamming andmore capable in deep attenuation environments.

• Less radiation environment -> enables commercial of the shelf (COTS)components.

• However, a major drawback due to the lower orbital altitude is the lackof coverage as indicated by Figure 2.5a looking at the footprint of LEOin relation to MEO, which implies having to launch additional satellitesin LEO to achieve similar coverage as MEO.


2.4.4 Space environment in LEOThe space environment is a harsh place, as a system engineer, one has to con-sider several aspects which the satellite needs to be able to mitigate such asthe vacuum of space, the great thermal gradients and the zero-g environment.But most importantly one has to mitigate the radiation from space [24].

Radiation

A crucial design constraint when developing a space system is the naturalspace radiation environment and its effect on the electrical components in thesystem. Long and short term radiation effects such as Total Ionizing Dose(TID) and Single Event Effects (SEE) provide different challenges [24]. Eachmission is unique, expressing different system requirements, as a result of thenatural radiation environment, in orbit, time frame, duration, and spacecraftdesign [24].

To further emphasize the damage space radiation has on space systems,through TID which is the cumulative amount of radiation dosage, measured ingray (Gy) and can affect transistor performance and through SSE created bysingle particles hitting the electronics and can affect the logic state of memoryor the output transistors through something called latch-up, which may causea high current state [16].

As earlier mentioned in section 2.4.3 LEO is preferable over higher al-titudes when it comes to radiation, since the dosage is much lower enablingCOTS components. It has been shown that at an altitude of about 500-600 kmin Sun Synchronous Orbit (SSO), see Section 2.4.6 for more about this specificorbit, at an average solar cycle the TID for over 2 years operations is about 30Gy if a 3 mm aluminium shielding is used around boxes, structures and arrays[10].

2.4.5 Orbital perturbationsThis section should explain what perturbations exist in LEO. The equationscan for now be ignored, but are listed here for future use when simulatingde-orbit time or the constellation’s change of coverage w.r.t time.

There exist several orbital perturbations that will affect the orbital decay ofa satellite such as aerodynamic drag, gravity anomalies due to the oblatenesseffect (J2) and solar radiation pressure [18].

Perturbations from a third body such as the moon will not be considered.Moreover, the orbits are restricted to be circular if no perturbing force would


act on them.In order to estimate how the semi major axis a of the satellite changes with

time as it orbits the Earth one applies the general equations of motion of asatellite relative to a non rotating geocentric equatorial reference frame andintroduces some perturbing force f to act on the satellite [18], see equation2.16,

d2r

dt2+

µ

r3r = f . (2.16)

Here r is the position vector of the satellite, µ is the gravitational parameterof the Earth and f is the acceleration due to all the perturbing forces on thesatellite [18]. One gets after some manipulation and differentiation to time thechange of a w.r.t time by equation 2.17, as follows

da

dt=

2a2

µV f , (2.17)

where V = drdt

is the velocity vector of the satellite.

Aerodynamic drag

The main perturbing force in LEO is expressed by the aerodynamic drag.Therefore a simplification to estimate a satellite’s life time can be done onlyconsidering aerodynamic drag. The decay lifetime is given by expressing theperturbing force as the aerodynamic drag as follows, see equation 2.18,

f = −CD1

2ϕA

MV · V , (2.18)

here CD represents the drag coefficient, ϕ the density of the atmosphere [18].One can distinguish the ballistic coefficient B as given by 2.19,

B =CD A

M. (2.19)

Non-Spherical Earth

Perturbations due to J2 effects of the gravity field also called oblateness of theEarth are active in the X-,Y- and Z-directions as follows, see equations 2.20,2.21 and 2.22

fx = −3

2µJ2

RE2

r5x(1− 5

z2

r2), (2.20)

fy = −3

2µJ2

RE2

r5y(1− 5

z2

r2), (2.21)


fz = −3

2µJ2

RE2

r5z(1− 5

z2

r2), (2.22)

where J2 is the second-degree zonal harmonic coefficient andRE is the Earth’smean radius [18].

Solar radiation Pressure

Perturbations due to solar radiation pressure can be expressed by equation2.23, as follows

f = CRW · AM · es

, (2.23)

where W is the power density of the solar radiation, M is the mass of thesatellite, c is the speed of light (in vacuum), A is the effective cross-sectionalarea of the satellite, Cr is the satellite’s reflectivity and es is a unit vector fromthe satellite to the Sun [18].

2.4.6 Possible orbitsMost Earth observation satellites orbit the Earth in near-circular orbits at anorbital inclination of 70 − 110, this allows a world coverage and a high res-olution of the observations [18]. Three orbits of interest for a LEO satelliteconstellation are:

• SSO

• Polar Orbit (PO)

• Equatorial Orbit (EO) as an extra orbital plane for further coverage ofthe equatorial region, where vessel density is high, see Figure 1.1.

To begin with, polar orbits and equatorial orbits are obviously at an incli-nation of i = 90 and i = 0.

Furthermore, the sun synchronous orbit takes place at a certain inclinationfor a given orbital height such that the node of lines rotates at the same rateand direction as the Sun in its apparent motion about the Earth [18] or simplerexplained: the orbital plane of the satellite rotates in space at the same rate asEarth moves around the Sun. This rate is around one degree eastwards eachday [22].

It is possible to calculate an SSO’s inclination considering circular orbitsby equation 2.24, as follows

cos(i)SS ≈ −0.098916(r

RE

)72 ↔ iSS = arccos (−0.098916(

r

RE

)72 ),

(2.24)


as can be seen it is dependent on the magnitude of the position vector of thesatellite, the distance between the center of Earth and the center of the satellite,r = H +RE [18].

Furthermore, to clarify in an SSO the Earth-Sun vector remains more orless perpendicular to the line of nodes during the entire flight of the satellite inLEO [18]. And if the orbit also is chosen as nearly polar the declination of theSun varies between 23, 5−−23, 5, at such conditions one can if launching atthe right time design an orbit that stay in sunlight for several months, creatingsomething called a SSO dawn/dusk orbit [18]. This could be desirable if onechooses to utilise solar power for power generation. Taking advantage of nearlyconstant illumination, but also that the solar arrays may be fixed relative to thebody of the spacecraft since the orbital plane is fixed relative to the solar vector,which gives a possibility of keeping the same attitude through the entire orbitwhile maintaining optimal power generation [22].

2.5 Satellite communicationsA short explanation to the equations used to produce a link budget and adata budget, see subsection 6.4.2. Have in mind that atmospheric losses andtransmitter- and receiver losses are not taken into account for the link budget.Whereas for the data budget, the satellite is considered to be passing straightabove a single vessel and the code rate is not taken into account.

To begin with, one can define the gain, G, of an ideal antenna in decibelsby equation 2.25, as follows

[G] = 10 log(4π · Ae

l2), (2.25)

where Ae is the physical aperture area, also known as the receiving cross sec-tion of the antenna, and l is the radio wavelength [14].

To perform a link budget, one desires to calculate the power of the receiv-ing antenna. As a starting point, one may assume that the power and gain ofthe transmitting antenna are known, Pt and Gt respectively, then the EffectiveIsotropic Radiated Power (EIRP), EIRP , can be defined by equation 2.26 asfollows,

[EIRP ] = [Pt] + [Gt], (2.26)

given here in decibel-watt (dBW). It is usually referred to as a figure of meritfor the transmission part of the linking [14].

We also need to calculate the free space path loss, Lfs, which obviously isdependent on the slant range, see again equation 2.14 and thus the minimum


elevation angle. Moreover, we can find the radio wavelength by consideringthe carrier frequency, f0, of the signal giving the radio wavelength by the wellknown formula

l =c

f0, (2.27)

where c of course corresponds to the speed of light in vacuum. Using equation2.27 and the slant range D, the free space path loss was calculated in decibels(dB) by equation 2.28,

[Lfs] = 20 log(π ·Dl

). (2.28)

Having found the free space path loss, Lfs, the EIRP, EIRP , by equations2.28 - 2.26 and considering circular polarization loss, Lp, contributing to a 3dB further loss, according to [20], the receiver power Pr could be calculatedby

[Pr] = [EIRP ] + [Gr]− [Lfs]− [Lp]. (2.29)

This, equation 2.29, is commonly known as the Link power budget equationand is from where the design and performance of the satellite link begins [14].

If we instead consider [Pr = Pr − Gr], one can define the sensitivity ofthe receiving station, Gr

T, [25]. As was the case of the EIRP, this can be seen

as a figure of merit for the receiving part of the linking [14]. It corresponds to

[Gr

T] = [Gr]− [Ts], (2.30)

where Gr is the gain of the receiving antenna and Ts is the system noise tem-perature at the Low Noise Amplifier (LNA). Using the sensitivity of the re-ceiving station, see equation 2.30, one can define the carrier to noise densityratio, [ C

N0], as follows

[C

N0

] = [Pr] + [Gr

T]− 10 · log(k), (2.31)

here k is the Boltzmann’s constant. The higher the value for the [ CN0

] the betterthe performance of a satellite communication’s link [14].

Finally, we can define the energy per bit to noise power spectral density,[Eb

N0], by adding the bit rate, BR, in decibel to equation 2.31 giving equation

2.32, as follows

[Eb

N0

] = [Pr] + [Gr

T]− 10 · log(k)− 10 · log(BR). (2.32)


It is interchangeable to the, [ CN0

], but is a better measurement for digital com-munications link performance [14].

Looking now instead to the data amount that can be sent during a satellitepass over a single subject. The time of contact, Tc, with an observer locatedat the surface of Earth, can be calculated using well known parameters fromearlier, see subsections 2.4.1 and 2.4.2, as follows

Tc =2 · P

2π · (Re+H)· Sw

2. (2.33)

As seen from equation 2.33, Tc is dependent on the Swath width, Sw, whichin turn is dependent on the minimum elevation angle, ϵ.

Using the time of contact, the data amount, DA, being sent to the satellitefrom a subject can be calculated by equation 2.34 as follows

DA = BR · Tc. (2.34)

2.6 Related work

2.6.1 Space NorwayAs of today Norway has an operational micro satellite program for maritimesurveillance with the satellites AISsat-1 and 2, NorSat-1 and 2 currently inorbit. Most interesting is NorSat-2 which was the first satellite with a VDESpayload in orbit, launched the 14th of July in 2017 [26]. A figure of NorSat-2in operational mode can be seen in Figure 2.6.

The Norwegian maritime surveillance program was founded by the Nor-wegian Coastal administration (NCA) and the Norwegian Space Centre (NSC)to assure safety and protection in Norwegian waters. The initiative was takenby the Norwegian Defence Research Establishment (Forsvarets Forskningsin-stitut - FFI) already in 2006 on basis of collecting AIS signals from vesselsindicating their position, identification and track [26].

A consortium was created where the satellite provider is the Space FlightLaboratory (SFL) at the University of Toronto Institute for Aerospace Stud-ies (UTIAS), which also is responsible for choosing a launch provider. Thepayload provider is Kongsberg Maritime and the satellite operations are per-formed by the NSC at their operation center Statsat [27].


Figure 2.6: Space Norway’s latest satellite NorSat-2, which carries a VDESand AIS payload, in ground tracking (limb pointing) operation, figure wasretrieved from [8].

2.6.2 ExactEarthExactEarth is the world’s leader in satellite AIS data services providing globalshipping and tracking capability, founded in 2009 and based in Cambridge,Ontario,Canada [28]. The company has recently finished a constellation of 67 satel-lites, providing a customer data latency as well as global revisit time of lessthan one minute [7], [2].

It is actually comprised of two constellations, where ExactView with its 9CubeSats provided by UTIAS SFL and others operating at 650 km altitude inequatorial and polar orbits is one [9] and the other is made up of 58 satellitescarrying on-orbit reconfigurable and reprogrammable payloads called App-STAR, an SDR-platform provided by Harris to give the capability of AIS,ASM and perhaps VDES in the future, on the Iridium constellation [29]. TheIridium constellation consists of a total of 75 satellites, whereas 9 are spares,the last satellites were launched as late as January 2019 to complete it [30].

The satellites in the Iridium constellation were built by the satellite providerThales Alenia Space and operate on a high LEO altitude of 780 km with an in-clination of 86.4 in 6 orbital planes, these are large and heavy (860 kg) satel-lites utilizing ISL to decrease latency [31]. One satellite links to two otherwithin its orbital plane (in-front and behind) and with two other in adjacentplanes (one on each side of the satellite) as can be seen in Figure 2.7. As a con-


Figure 2.7: The Iridium constellation in circular near polar orbit at 780 km al-titude, it also shows the communication interconnections between the satellitescalled Inter Satellite Links (ISLs), figure was retrieved from [32].

sequence the constellation has always constant communication with multipleground stations, hence messages can be received and delivered in ”real-time”according to Kocak and Browning [29].

Moreover, Proud, Browning, and Kocak [2] adds that ExactEarth con-tracted Kongsberg Maritime and SRT Marine to provide AIS payloads fordetection of AIS class A and class B vessels in the ExactView constellation.

Chapter 3

System engineering, stakehold-ers and use-cases

3.1 A system engineering approachIn order to fully understand the whole system and to work in a concise andpractical way the author has chosen initially to follow the system engineeringapproach described by Larson [5]. It is an iterative process, steps are donein order but as the system develops new information come to light that mightchange earlier steps in the process. As a starting point the system is definedby the needs of the stakeholders. This is done by identifying the stakeholdersto analyse on one hand how they influence the system and on the other handhow the system influences them.

Moreover, a trade-off was made to determine the key stakeholders and fromthem the stakeholder requirements were established to pinpoint the mission re-quirements, known as well as the critical acceptance criteria. Use-case scenar-ios were analysed and together with the stakeholder - and mission requirementslay the foundation to form the system requirements. From this and regardingthe work done by Øystein et al. [10] system requirements could be synthesizedusing the approach laid out by NASA [11].

Furthermore, a work breakdown structure was developed to formulate howthe system can be broken down into compositions of the main system. Havingdone this a simple system architecture was established and a Functional FlowBlock Diagram (FFBD) was made, showcasing the development and disposalof a typical satellite in the system.

Taking into consideration the system requirements and the system archi-tecture, the design was proposed by developing budgets regarding mass, size,

28

CHAPTER 3. SYSTEM ENGINEERING, STAKEHOLDERS ANDUSE-CASES 29

power and linking following the approach laid out by Wertz and Larson [12]in Space Mission and Design Analysis. Lastly, a risk assessment has beencarried out to identify potential risks with the development, operations anddisposal of the system.

3.2 StakeholdersThe reason one develops a system is to meet a need from customers and otherstakeholders. A stakeholder being defined as ”a stakeholder is a person or anorganization that has an interest in, or can be affected by, the system or project”[5]. This is because the stakeholders have an interest in the need, the mission,or the operational capability of the system.

Moreover stakeholders can be organized into passive - and active stake-holders as well as sponsors. An active stakeholder actively interacts with thesystem when operational and in use, on the other hand a passive stakeholderinfluences the success of the system, but doesn’t actively interact with the sys-tem. In addition there is the sponsors who buy the product or fund it with aspecific mission in mind. Sponsors can be both passive and active stakeholders[5].

The identified stakeholders and their classifications are presented in Table3.1.

3.2.1 Stakeholder interestTo further explain the stakeholders and their interest in the system, a morethorough explanation is presented here.

Maritime authorities

”The European Maritime Safety Agency (EMSA) is one of the EU’s decen-tralised agencies. Based in Lisbon, the agency provides technical assistanceand support to the European Commission and Member States in the develop-ment and implementation of EU legislation on maritime safety, pollution byships and maritime security” [33]. It supports ESA in its pursuit of develop-ing satellite VDES activities within the frame of the ARTES programme [34],aimed at identifying potential satellite VDES benefits through collection ofneeds, requirements and experience of potential users [27].

”The International Association of marine aids to navigation and Light-house Authorities (IALA) is a non-profit, international technical association.

30 CHAPTER 3. SYSTEM ENGINEERING, STAKEHOLDERS ANDUSE-CASES

Table 3.1: Various stakeholders in the satellite VDES system categorized asdefined by Larson [5].

Stakeholder Classification Description

Customers/Users Active, SponsorInterested in whatthe system can provide.Pays for the system.

Satellite provider Active Provides the satellites.

Service provider Active, Sponsor

Provides the serviceof the system to customers/users.Invests in the developmentof the entire system.

Payload provider Active Provides the SDR-platform.

Satellite operator Active

Operates the satellitesin orbit. Appliescontrol and supportto the satellite constellation.

Ground stationprovider Active Provides the

ground station network.

Launch provider Active Provides the capabilityto go to space.

Separation systemprovider Passive

Ensures the separation ofsatellites from launch vehicle when in orbit.Further entitled to decreasethe mechanical loadson the satellites at launch.

Scientists ActiveInterested in developing thesystem in exchangeof investments in research.

Competitor Active

Other payload providers,influences the timeof development andcost of the system.

Space agencies Passive, SponsorProvides support andresearch for satellite VDEStesting, development, etcetera.

Maritimeauthorities Passive Mandates regulations for

the maritime industry.

Radio communicationauthorities Passive

Decides upon frequencyallocations for the system.Sets up a frameworkof requirements on the system.


Established in 1957, it gathers marine aids to navigation authorities, manufac-turers, consultants, and scientific and training institutes from all parts of theworld and offers them the opportunity to exchange and compare their experi-ences and achievements” [35].

IALA supports the development of VDES and is also the leading organi-sation in creating a framework for the ongoing research.

”The International Maritime Organization (IMO) is the United Nationsspecialized agency with responsibility for the safety and security of shippingand the prevention of marine and atmospheric pollution by ships” [36]. Theorganization is responsible for the endorsement and publishing of the interna-tional convention for the safety of life at sea (SOLAS), the most important ofall international conventions dealing with life at sea [37].

IMO restricts all class A vessels, such as vessels of 300 gross tons andupwards engaged on international voyages, cargo ships of 500 gross tonnageand upwards not engaged on international voyages and most importantly allpassenger ships irrespective of size, to carry AIS according to the SOLASagreement chapter V: ”Safety of navigation” [37].

A further decision, making it mandatory to change from AIS to VDES forvessels as described above has to be made by IMO to not delay the operation ofboth the terrestrial VDES system and the satellite VDES system several years[38].

Radio communication authorities

”The International Electrotechnical Commission (IEC) is a not-for-profit, quasi-governmental organization, founded in 1906. The IEC’s members are NationalCommittees, and they appoint experts and delegates coming from industry,government bodies, associations and academia to participate in the technicaland conformity assessment work of the IEC” [39].

They influence the system by the test houses and standards that need to befinalized, in order to know what requirements are put on the system.

The International Telecommunication Union (ITU) is the United Nationsspecialized agency for information and communication technologies. As de-scribed on their website ”founded in 1865 to facilitate international connectiv-ity in communications networks, we allocate global radio spectrum and satel-lite orbits, develop the technical standards that ensure networks and technolo-gies seamlessly interconnect, and strive to improve access to ICTs to underserved communities worldwide” [40].

ITU invites the private sector and universities to the WRC in November


2019, where a decision regarding frequency allocation of frequencies 156-163MHz for downlink of satellite data has to be taken for further development ofthe satellite VDES system [38].

Customers

Interested in what the system can provide, satellite data (ship position, speed,route, etcetera). As the customer pays for the system they regulate the priceof the system by what they are interested in paying. According to Lázaro etal. [1] the system shouldn’t be more expensive for the customers than the AISsystem in use today. The customers are usually also users of the system, butdon’t necessarily have to be.

Possible customers/users: mariners, coast guards, port authorities, nationalmaritime authorities, naval forces, etcetera.

Satellite provider

The satellite provider provides the satellites and influences the space segmentof the system by several constraints such as mass, power, size, temperaturecontrol, communication system and attitude and orbital control.

Possible providers: GomSpace [41], ÅAC Clyde Space [42], Blue CanyonTechnologies (BCT) [43], Innovative Solutions In Space (ISIS) [44] and Uni-versity of Toronto for Aerospace Studies (UTIAS) Space Flight Laboratory(SFL) [45].

SFL provides experimental satellites, in specific interest of this study, theybuilt the satellites NorSat-1 and NorSat-2 in cooperation with the NSC carry-ing AIS- and VDES payloads [26]. It has also provided satellite platforms forExactEarth’s constellation Exactview, as for example satellite EV-9 flying inan equatorial orbit to pick up AIS signals from vessels around the equator [9].

Service provider

The party that provides the service of the system to the customer/user. Onecould argue that it is the stakeholder that invests in the system in order to sellthe service to the customers/users. This is done by collecting data from thesystem in order to enhance it and provide it through a desirable interface, suchas for example a web interface where customers/users can follow their vesselscontinuously anywhere on the globe. The service provider could be the sameparty as the satellite operator and/or satellite provider, but doesn’t necessaryhave to be.


Several potential service providers will be listed later in Table 5.1, but a fewexamples are ExactEarth, as earlier discussed in Section 2.6.2, ORBCOMMwhich claims to be the global market leader in terrestrial and satellite AISdata services [46] and lastly a relatively new company called Spire Global,founded in 2012, that offers among other things vessel detection by AIS fromits satellite constellation, the largest satellite AIS constellation at the time ofwriting, see again Table 5.1, which they also operate themselves. In addition,the satellites and payloads are designed, built and integrated in house [47]. Inthis case they are not just a service provider, but also the payload provider,satellite provider, satellite operator and ground station provider [47].

Payload provider

Provides the SDR platform which comes with constraints on the satellite toprovide power, structural dimension and orientation of antenna. Further re-quirements on the system to have sufficient signal strength and coverage. De-sign to cost, production and installation efficiency.

Possible provider: Saab TransponderTech, taking into consideration thatthis report is done on request from Saab, the author considers the companyto be essential for the continuance of the project. As of writing this thesis,the airborne transponder R5A [48], used on helicopters and Unmanned AerialVehicles (UAVs), is envisioned to be developed into a payload for satelliteVDES.

Launch provider

Provides the capability to go to space, influences the system by limitations onhow many satellites that can be launched at the same time as well as what alti-tude and inclination is possible. Moreover, one has to consider the vibrationalloads at launch. The launch provider also chooses the launch location.

Possible providers: SpaceX an American company that provides launcheswith a falcon-9 or falcon heavy launch vehicle. Indian Space Research Or-ganization ISRO provides launches with the Polar Satellite Launch Vehicle(PSLV). Blue origin, another American company, provides launches with thenew Shepherd launch vehicle. Astrium is a French-German company provid-ing launches with Ariane-5. Energia is a Russian company providing launcheswith a Soyuz launch vehicle. Lastly, Rocket Lab who offers launches with itslaunch vehicle Electron specifically adapted for launching small satellites.

To mention a few, a whole list is provided in the report ”State of the Art:Small Spacecraft Technology” provided by NASA [16].


Separation system provider

Ensures safe separation of satellites from launch vehicle when in orbit. De-creases the mechanical loads on the satellites at launch.

Possible providers: Spaceflight industries with their P-Pod, University ofTokyo with their T-pod, UTIAS SFL with their X-pod, ISIS with their ISIPOD,Ecliptic enterprises with their Rocket Pod and Planetary systems corporationwith their Canisterized Satellite Dispenser (CSD), to mention a few [16].

Satellite operator

Operates the satellites in orbit. Applies control and support to the satelliteconstellation. This is done by tracking, transmitting commands and receivingtelemetry data through the ground station network to/from the satellites.

Possible providers: usually the same as the satellite providers or the groundstation provider.

Ground station provider

Provides the ground station network, functioning as a communication inter-face between the satellite constellation in space and the satellite operations andusers/customers on ground. Provides downlink and uplink capabilities to thesatellite constellation (telemetry, tracking, command and VDES messages).Possible providers: Swedish Space Centre (SSC) and Kongsberg Satellite Ser-vice (KSAT).

SSC’s global network of ground stations was specifically designed to pro-vide comprehensive communications and ground support to Earth-orbitingsatellites [49]. Nonetheless, it is unclear how many actual sites that SSC canprovide.

KSAT owned by Space Norway and Kongsberg Defence and AerospaceAS provide 170 remotely controlled antennas with more tan 20 sites world-wide. Aims to provide optimized locations for satellites in polar, inclined andequatorial orbits [50].

Space Agencies

Mainly the European Space agency (ESA). Provides support and research forsatellite VDES testing, development, etcetera. This is done within the al-ready mentioned ARTES programme, which is a programme that supportsresearch and development within satellite communication applications to pro-vide worldwide leading applications [34].


Competitors

The competition works as a benchmark for the development of the system.Possible competitors: From the position of the author those are mainly pay-

load providers such as Kongsberg Maritime [51], GomSpace [52] and Harris[53].

Kongsberg Maritime develops similar products as Saab, but whereas SaabTransponderTech has reached far when it comes to the terrestrial part, Kongs-berg Maritime in cooperation with the NSC are further at the space segment,already testing experimental satellite VDES in satellite Norsat-2. KongsbergMaritime has also had tests done on-board the International Space Station(ISS) and on board satellites in ExactEarth´s constellation ExactView [26].

Scientists

Different research facilities like universities, can provide knowledge and workpower in exchange of investments and education. Can be a great contributionto the development of the system.

Possible providers: KTH Royal Institute of Technology, Linköping Uni-versity and Chalmers University, to mention a few that are close to Saab loca-tion wise.

3.2.2 Identifying key stakeholdersTo identify the key stakeholders a trade-off has been done regarding the keytraits: impact, replace-ability, identifiability and relationship dynamics ac-cording to the method presented in [54]. The result can be seen in Table3.2, the key stakeholders were identified as the customers/users, the satelliteprovider, the payload provider, the satellite operator and the service provider.

To describe the key traits:

• Impact: How crucial is the particular stakeholder for the system? Is theimpact very high for financial, service or knowledge reasons?

• Replace-ability (Re-Ab): Takes into account how dependent the sys-tem is upon the stakeholder. Can the system proceed without the stake-holder?

• Identifiability (Id-Ab): Is the stakeholder clearly identifiable in the sys-tem? Is it straight-forward to see what role it has in the system?


Table 3.2: Trade-off table to identify key stakeholders.

Stakeholder Impact Re-Ab Id-Ab Re-Dy Result

Customers 1 1 1 1 4

Satelliteprovider 1 0 1 1 3

Payloadprovider 1 1 1 1 4

Satelliteoperator 1 0 1 1 3

Serviceprovider 1 1 0 1 3

Launchprovider 1 -1 1 -1 0

Separationsystem

provider0 -1 1 -1 -1

Scientists 0 -1 -1 1 -1

Competitors 0 0 1 0 1

Spaceagencies 0 1 0 1 2

Maritimeauthorities 1 1 1 -1 2

Radiocommunication

authorities1 1 1 -1 2

Legend Weighting:

1: High FactorCriteriaweightedequally

0: Neutral factor-1: Low factor


• Relationship dynamics (Re-Dy): Is the relationship with the system dy-namic or static? Is the relationship growing as the system proceeds?

The grading system is done according to the following criteria: To begin with,a minus one (-1) represents that the impact of the stakeholder is low enoughto not jeopardize the mission, the stakeholder is replaceable meaning that itsfunction can be taken over by another stakeholder, it isn’t possible to make aclear identification of what the stakeholder is providing to the system and therelationship dynamics is static and is not growing with time.

A zero (0) represents a moderate level of impact, replace-ability, identifi-ability and relationship dynamics.

Lastly, a one (1) represents that the stakeholder has a high impact on thesuccess of the mission, that the stakeholder is not replaceable, that one clearlycan identify what the stakeholder is providing to the system and that the rela-tionship dynamics is dynamical with the system and set to grow over time.

3.3 Context diagramHaving described the different stakeholders in more detail, one can also lookat how the stakeholders interact with the system as can be seen in Figure 3.1.


Satellite operator

Satellite VDES system

Payload providerScientists

Customers

Satellite provider

Launch provider

Active stakeholders

IMO

IALA

ITU

IEC

Passive stakeholders

ESA

EMSA

CompetitorsGround station

provider

Supports development of Satellite VDES

Supports development

of Satellite VDES

Regulations:need mandatory

decision for Class A vessels to change from AIS to VDES

Invites to WRC:Where decision

regarding frequency allocation 156-163 MHZ for SAT-downlink is to

be taken

Offers exchange of information between authorities,

influences decisions taken at WRC

Imposes standards and

test houses

service

needs

REQ on satellite

REQ on payload

Telemetry and

trackingCommand

Launch capabilities

Adapt system for

compatibility

Seperation System Provider

Seperation capabil ities

Data distribution

REQ on ground station network

Service provider

investments, enhancement of

dataData

Figure 3.1: A context diagram of the stakeholders and how they affect and areaffected by the system.

3.4 Use-casesThere are several use-cases that have to be considered for a satellite VDES sys-tem. These have been established in regard to where a need for over the horizoncapability to reach vessels at open water has been identified. Lázaro et al. [1]have identified four different use-cases for the satellite VDES system: SAR,broadcast of Maritime Safety Information (MSI), transmission of Facilitationof International Maritime Traffic (FAL) forms and download of updated dig-ital publications. Hence these use-cases have been chosen to be investigatedfurther. The interaction of the system’s segments for each regarded use-casecan bee seen in Figure 3.2. Note that Vessel Traffic Service (VTS) regardscommunication between ships and port authorities to guarantee navigationalsafety when a vessel enters a port, this use-case has not been explained furthersince it is not part of the satellite segment of the VDES system, but insteadillustrated here as a typical use-case between the shore and the vessel seg-ment of the whole VDES system, see the exchange between vessel 3 and theuser. Another similar example is the Sea Traffic Management (STM) routeexchange between Vessel 1 and Vessel 2, which is within the ship segment of


the terrestrial VDES.

SAR

The IMO convention SAR service allows vessels to communicate for searchand rescue purposes. A satellite VDES system would allow vessels in need ofrescue at open sea to transmit through (uplink to satellite) the system informingnational maritime authorities at shore.

Download of updated digital publications

An example could be the download of the Global Maritime Distress and SafetySystem (GMDSS) master plan containing contact information of the nationalauthorities of all countries. This would be sent out from the shore segmentthrough the satellite segment and downlinked to the ship segment.

Transmission of FAL

The IMO convention Facilitation of international maritime traffic forms aresent from ships to national maritime authorities in order to provide informationconcerning for example type of cargo, number of crew members and passengerlists. Hence a satellite VDES system could increase the speed of administra-tion when entering national waters of a country.

Broadcast of MSI

To support national maritime authorities to transmit MSI regarding possibleareas to be avoided or perhaps weather alerts due to extreme weather. Thiswould also take place from shore segment over the horizon through the satellitesegment and downlinked to the ship segment.

Route exchange

This part is not envisioned by Lázaro et al. [1], nevertheless this use-case willbe analysed further keeping in mind the benefits of route exchange (as deter-mined by the STM validation project [55]) for broadcasting a vessel’s position,velocity and intended route by having global coverage with low latency delay(<2 s) which a continuous global satellite constellation is desired to offer.

This use-case could for example be provided by uplink from vessel to satel-lite regarding the vessel’s position, route and speed and then downlinked tonearest ground station in order to send the information to a data center where


the service provider could enhance the data to provide a web interface for itscustomers/users to be able to follow their ships and other ships of interest atopen waters continuously. This use-case scenario can also be seen in Figure3.2.

SAR

GMDSS

STM route

exchange

FAL

Satellite operator

ISL

User

MSI

TT & C

IP

Vessel 1

Satellite 1Satellite 2

Payload data

Ground station

Service provider

Web interface

Data center

MCC

Vessel 3

VTS

STM route

exchange

Vessel 2

Figure 3.2: Potential use-cases when the system will be operational and howthose interact with the user, the satellite operator and the service provider.Pay attention to how the ground-station network uses IP to connect with theMission Control Center (MCC), the data center and possibly directly to theuser.

3.5 System limitationsFirst of all, to define the system requirements one has to consider the limi-tations of the system and the dependencies. To see where these constraintsmatter a Work Breakdown Structure (WBS) has been carried out for the sys-tem, as can be seen in Figure 3.3.


SDR payload

Satellite VDES

Satellite platform

SW/FW Antenna

Constellation

AOCS EPS

ASMVDE AIS

Power generation

Energy storage

Hardware

Structures and

Mechanisms

Commun-ications

Command and Data Handling

Thermal controlsystem

CubeSat standard

On-Board Processing

House keeping

computer

Determination and control computer

Payload on board computer

Altitude Coverage

Multipath effect

Power distribution

Number of satellites

Feed link

Swath width

Electromagnetic radiation

Orbit perturbations

Launch segment

Inter satellite

link

Ground segment

Attitude control

GNSS

Attitude determination

Data Latency

delay

Signal strength

Temperature gradients

Batteries

Solar panels

Satellite availability Modulation

and coding scheme

Sun pointing

Power regulation

and control

Uplink and

downlink

Uplink Uplink

Nadir pointing

MCCGround station

networkLaunch vehicle

Seperation system

Passive thermal control

Space segment

User link

Deployment mechanisms

Orbit determination

Inter satellite pointing

Inclination

Minimum elevation

angle

Power generation

Symmetry

Relative latitude

coverage

Figure 3.3: A WBS of the system.

Limitations

• Satellite availability - How often is a satellite available at an arbitrarypoint at the surface of Earth?

• Satellite noise level - How is the VDES payload impacted by other satel-lite subsystems interfering with its operational frequency band?

• Data packet collisions - How does the satellite handle thousands of in-stantaneously received data packages from ships in its FOV if enteringa condense area?

• Orbital propagation - solar radiation pressure, aerodynamic drag, earthoblateness effect - How does perturbations disturb the desired orbit?

• Detection rate - How high is the probability that the satellite detect eachindividual ship in its FOV?

• Signal strength- What signal strength is attainable from the decided or-bital altitude?


• Power - What power constraints exist due to restrictions from the satel-lite platform?

• Mass - How to minimize mass considering the high launch costs?

• Temperature - What temperature differences can be expected consider-ing the operational environment?

• Dimension - What restrictions exist dimension-wise for the satellite?

• Uplink and downlink - How is communication transmissions handled,what constraints exist?

• Inter Satellite Link (ISL), how does it affect the possible distance be-tween satellites within the constellation considering the required an-tenna gain and power need for transmission in relation to distance?

Dependencies

As recognized by Eriksen et al. [27] transmission from satellites to ground isdependent on:

• The distance between satellite and ground target also known as slantrange (in turn dependent on orbital altitude and elevation angle).

• The multipath effect - diffuse scattering of the microwave on the seasurface.

• The ionosphere of Earth.

Other dependencies are described by Reid et al. [17] who take a look at thelatency delay of a global constellation. Latency delay is a measurement onhow long it takes for a user having requested information from the system toreceive requested information. It is dependent on:

• Incorporating ISL between satellites in the constellation. How muchcan this reduce the latency delay?

• Attainable global coverage (dependent on orbital altitude, number ofsatellites).

Chapter 4

System design and architecture

Following the approach laid out by Wertz and Larson [12], see Figure 4.1, adecision table was carried out in order to evaluate what was optional in thesystem design and what had already been set by the stakeholders, use cases,system requirements and the WBS, see Tables 4.1 - 4.2.

Figure 4.1: The concept characterization process that this work has followed,where the subject has been defined by the vessel segment of the VDES systemas explained in section 2.1, the figure has been retrieved from [12].

43

44 CHAPTER 4. SYSTEM DESIGN AND ARCHITECTURE

Table 4.1: Emphasizes the design choices that were set and the ones whereoptions were available.

Element Can be traded Reason

Mission concept No

SAT-VDES: defined as communication and trackingwith/of vessels at sea by satellite,

using VDES inthe maritime band.

Subject No

Vessels at open seacarrying AIS/VDES transponders

out of sight ofshore segment, well defined.

Payload No

SDR-platform from Saab.Frequencies and hardware are

set, software andstructure are changeable.

Spacecraft Bus PartlyMultiple options, however

restricted by theCubeSat standard.

Launch vehicle Yes Choose minimum costfor selected orbit.

Separation system PartlyChoose minimum cost,

restricted by CubeSat standardand launch vehicle.

Orbit Partly

Specified to be incircular LEO 400-700 km

to minimize costs andrequired orbit maintenance,

however, inclination is not set.

Ground system Yes

Use existing ground stationnetwork OR

construct new dedicatedground station network OR

utilize a combinationof base stations by Saab

and a fewer numberof ground stations.

Communicationsarchitecture Partly

User link:fixed by subjectrestrictions on

radio frequency 156-163 MHz.Feed link:

flexible choice of radio frequency.ISL: high throughput radio frequency

link or optical link.

Mission operations Yes Adjustable level of automation

CHAPTER 4. SYSTEM DESIGN AND ARCHITECTURE 45

Table 4.2: Emphasizes all areas of the WBS, see Figure 3.3, where a decisionregarding approach had to be evaluated.

Mission element Option Area SAT -VDES options

SAT -VDES decided

Mission concept Datadelivery

Distribution of data to users at shorethrough dedicated ground stations or

base stations or a combination ofground stations and base stations.

And distribution of datadirectly to subjects.

Down - and up linkto users and ground stations.

Tasking Autonomous tasking.

Subject What isto be sensed

Vessels carryingVDES/AIS transponders.

Payload Frequency Radio VHF:156-163 MHz.Size vs.

sensitivitySmall apparatus with high

power.

Complexity Single or multiple antennas. Single band: The maritimeband.

Spacecraft bus Propulsionsystem Not desired.

Orbit control On-board control.Navigation: on-board GNSS.

ADCS

3-axis control.Payload pointing: nadir pointing.Spacecraft pointing: sun pointing,

inter satellite pointingand nadir pointing.

Structure 3 U - 6 U Restricted to CubeSat standard.

Power Solar panels: body mountedor deployable.

Power source: solar arrays.Power storage: batteries.

Launch system Launchvehicle

Determined by orbit,spacecraft structure and cost.

Separationsystem Optional dispenser. Restricted to

CubeSat standard.launch

siteDetermined

by launch vehicle.

Orbit Specialorbits Circular LEO orbit

Altitude 400-700 km. Common altitude for allsatellites.

Inclination 0°- EO or 90° - PO or 88-98°- SSO Common inclination for allsatellites.

Constellationconfiguration

Walker delta geometry or Walker star geometryor other geometries.

Minimize number of satellitesfor global continuous coverage.

Ground system Existingor dedicated

Ground network provideror build dedicated ground system.

Communicationsarchitecture Timeliness Real

time link.Control

anddata dissemination

Data dissemination - direct to userand/or to data center. Control - MCC

RelayMechanisms

Radio frequency for non-userlink not decided yet.

Three links:ISL, feed link and user link.

Missionoperations

Automationlevel

Fullyautomated satellite control

or part time operated.Autonomy

levelPartial autonomy or full

autonomy.


4.1 System segmentsAs earlier mentioned, the system is composed of three main segments, seefigure 3.3. This is further showcased in Figure 4.2.

• The Space segment: includes the satellites carrying payloads of SDRsfor VDES purposes in a constellation designed to bring continuous globalcoverage.

• The Ground segment: includes the MCC, the ground station networkand the data centre. It is an interface between the space segment and theuser segment.

• The Launch Segment: includes the launch vehicle and the separationsystem. It offers access to space to the the space segment.

Space segment

Ground

segment

Data center

Satellite

Launch

vehicle

Launch

segment

Ground

station

Satellite Satellite Satellite

Seperation

system

ISL ISL ISL

Feed

link

User

segmentUser

link

IP

Vessel User

MCC

Figure 4.2: A simplified architecture of the system where the yellow arrowrepresent the user link at VHF, the blue arrows represent ISLs and the redarrow the feed link. The black arrows represent ground network (IP) links.

Moreover, the user segment, corresponding to vessels at sea and users at shore,is also given in Figure 4.2 since it interacts with the system, however, it is notdefined as part of the system.


4.2 Functional architectureDevelopment of a typical satellite in such a constellation could be explainedby a Functional Flow Block Diagram (FFBD) as shown in Figure 4.3.

Integration of

payload(s) in

satellite bus

Integration of

satellite into

launch vehicle

LaunchDecision to

launch satellite

De-tumble

De-orbit End of life

Deploy CubeSat

into parking

orbit

Seperate

seperation

system from

launch vehicle

Continue

operations?

Positive decision for

frequency allocation of

satellite VDES at WRC

Development

of VDES

payload

Development

of satellite

bus

Testing

Calibrate

instruments

Production

Perform

nominal

operations

Perform off-

nominal

operations

OROR

Yes

No

Deploy antennas

and solar panels

Initiate

spacecraft

functions

Initiate payload

functions

Figure 4.3: The development of a typical satellite in a LEO constellation forVDES from decision at WRC to end of life and de-orbit phase.

4.3 Space segment

4.3.1 Satellite constellationA constellation is a set of satellites distributed over space intended to worktogether to achieve common objectives as for example optimizing coverage orgaining global coverage, if the satellites are flying closer together it is calleda cluster or formation [13]. To further envision this, one may take a look atFigure 4.4, which gives an idea of the need of control in relation to the distancebetween satellites.

In the interest of establishing a satellite constellation for Earth observationpurposes, several drivers have to be identified. As identified by George et al.


Figure 4.4: Gives an understanding of the need of control in relation to thedistance between satellites, figure was retrieved from [56].

[57] the main drivers are:

• Mission imposed requirements (for example acceptable latency delay)

• Degree of coverage

• Number of planes and launches

• Orbital altitude

• Cost

Nevertheless, one of the most important aspects to take into account whendesigning a satellite constellation is that, after decades, still no absolute rulesexist for constellation design [13]. Having said that, looking back at the maindrivers, the following constraints, as seen in Table 4.3, have been set:

Table 4.3: Constellation design drivers

Requirement IdentifiedMission imposed requirement One-way latency delay < 2 s

Degree of coverage 100%

Regarding cost, distinctively the number of satellites is the principal costdriver, hence it is desired to achieve continuous global coverage with the min-imum number of satellites possible. Higher altitudes offer more coverage asearlier discussed in subsection 2.4.3 and thus need less satellites to achievedesired coverage. However, since one may need larger and more complexsatellites for higher orbital altitudes due to for example tougher radiation envi-ronment, a required propulsion system in order to not stay longer than the limit


of 25 years, higher power required for transmission and a even further com-plex MCS when dealing with many more subjects within the satellite’s FOVto avoid data packet collisions from happening. It is not necessarily cheaperthan launching additional smaller satellites in a lower orbital altitude [13].

In this analysis the author has chosen to go for the second approach usingCubeSats in circular LEO, especially having considered that the second prin-cipal cost driver is the launch costs, which also represents the highest risks asonly 90 % of all launches are successful [13]. Bearing in mind that the higheraltitude the more complex satellite systems are required, which has a tendencyto increaser the launch costs due to the increased mass. As well as the obviousreason for a cost increase, one needs an increased delta V to reach a higheraltitude.

Another driver for constellation design is Earth’s oblateness effect thatmake constellations of different altitudes, eccentricity and inclination driftapart fast over time [13]. This is due to the fact that the oblateness of Earthrepresented by J2 causes the right ascension of the ascending node and theargument of perigee to rotate fast in relation to most satellite constellation’slife time. These rotation rates are a function of inclination, eccentricity andaltitude and hence vary with these parameters [13].

Therefore to maintain the constellation structure, in general constellationshave satellites at a common altitude, inclination and eccentricity as rotation isnot a problem if the whole constellation rotates together, a few exceptions existsuch as the addition of an equatorial orbit for which earth’s rotation doesn’tmatter [13]. Have in mind that in LEO not even propulsive manoeuvres canrealistically cancel these effects, due to the amount of fuel needed to stop thenode drift rate between satellites at different inclination, altitude or eccentricity[13].

An initial analysis has been made, where the conclusion that the cheap-est regarding launch costs can be achieved using a satellite constellation ofcommon altitude, eccentricity and inclination. An example of one such con-stellation is the Walker delta constellation which is defined by circular orbitswith the following parameters [13].

• i - Common inclination for all satellites

• T - Number of satellites

• P - Number of orbital planes evenly spaced in node

• F - Relative spacing between satellites in adjacent planes


Figure 4.5: A Walker star constellation of six orbital planes, as viewed whenlooking down on the north pole. One can see the counter rotation betweenplanes 1 and 6 and thus the nodal spacing is different than for the rest of theplanes, figure was retrieved from [22].

A walker delta constellation is therefore shortly described by i:T/P/F to specifyhow the constellation is ordered [13].

Another example is the Walker star geometry, illustrated here by Figure4.5, which is a special case of the the Walker delta geometry, likewise it hascircular orbits of common inclination and orbital altitude, with a certain num-ber of orbital planes. It is different however, as it is a constellation of polar(or near polar) inclination and due to risk of collisions cannot have all or-bital planes at the same nodal spacing due to counter rotating adjacent orbitalplanes.

Furthermore, as a polar constellation it is of special interest to this study asit simplifies the estimation for global coverage, see again theory on coverage2.4.2. An example of a Walker star constellation is the, earlier mentioned,Iridium constellation, see Figure 2.7.

4.3.2 PayloadPoghosyan and Golkar [15] see great potential in the CubeSat’s capabilityfor multipoint sensing and enhanced coverage, stating it as one of the mainstrength of the CubeSat platform. This they argue is made possible by thelow-cost and fast development cycle of the CubeSat platform, enabling excep-tional scientific missions based on distributed space system architectures suchas constellations.

One such mission is Earth applications from space, which Poghosyan andGolkar [15] argue is attracting increasing interest, they further mention ship


tracking by detecting AIS signals from space as one such application. This isof course of relevance to this work and further strengthens the argument thatsatellite VDES by a CubeSat constellation, building on the AIS space heritage,is a future area of interest to Saab.

The starting point of the envisioned payload is as mentioned earlier Saab’sairborne transponder R5A, see subsection 3.2.1. It is an airborne transponderbased on the latest SDR technology qualified for airborne standards such as theenvironmental conditions and test procedures for airborne equipment (RTCAD0-160G) [48]. It is currently adapted for AIS, but due to the flexible SDRplatform (see Section 2.3) an upgrade to VDES is not far from becoming areality. To give a better understanding of its current form factor (144 mm ×65 mm × 200 mm), see Figure 4.6.

Figure 4.6: Saab’s Airborne transponder R5A, figure was retrieved from [48].

The payload is according to Wertz and Larson [12] the most significantdriver of spacecraft design. They emphasize that the payload’s physical pa-rameters size, weight and power rules the physical parameters of the space-craft. Therefore, see Table 4.4, a design proposal regarding those physicalparameters is given for a future payload concept from Saab, slightly regulatedin order to fit a 3 U (single payload) or 6 U (one of several payloads) CubeSat.See Figure 4.7 to get a grasp of the CubeSat standard.


Table 4.4: The physical parameters for the envisioned payload from Saab forsatellite VDES, downscaled to fit the form factor of a 3 U CubeSat. Note:Orbit Average Power (OAP), Peak Power (PP).

Type SpecificMass 1 kgPower 20 W (OAP), 60 W (PP)Size 150× 98× 98 mm3

4.3.3 Satellite platformThe satellite platform also known as the satellite bus has several functions allrelated to its payload that can be summarized as follows [12]:

• Support the payload mass.

• Point the payload correctly.

• Keep the payload at the right temperature.

• Provide electrical power.

• Telemetry, Tracking and Command.

• Put the payload in the right orbit and keep it there.

• Provide data storage and communications.

The satellite platform is composed of 5 subsystems which aim to provide thesefunctions called Power system, Guidance Navigation and Control (GNC), Ther-mal control, Communications and Command and Data Handling (C & DH).

Limitations for the satellite platform have been decided having taken intoconsideration stakeholder requirement SH-3.4, see Table 6.1, in order to re-duce costs and development time of the satellite platform.

• Satellite platform in accordance to the standardized CubeSat criteria.

• No propulsion system.

• No active thermal system.

As the satellite platform had been decided to be in accordance to the Cube-Sat standard, here follows a small explanation of the standard.


The CubeSat project started out in 1999 as a collaborative effort betweenCalifornia Polytechnic State University (Cal Poly) and Stanford University’sSpace Systems Development Laboratory (SSDL). The original purpose wasto provide affordable access to space for universities [58]. It has turned outto be a major success with over a 100 universities, high schools and privatefirms developing pico/nano satellites carrying integrated scientific, private andgovernment payloads [59].

As of today, having expanded outside the university world, the primarymission is to provide access to space for small payloads [59].

A CubeSat is a small satellite that must conform to particular criteria re-garding size, shape and mass. The very specific criteria reduces costs by al-lowing companies to mass produce components and offer off the shelf parts.Moreover the standardization provides further cost reductions by flexible trans-portation and deployment in space [58].

CubeSats are defined by the CubeSat unit refereed to as 1 U defined as 10cm × 10 cm × 10 cm with a mass between 1 - 1.3 kg [58]. In recent yearsCubeSats have grown larger to provide greater CubeSat capabilities and arenow offered as 1.5 U 3 U, 6 U to mention a few [15], the CubeSat specificationcan be seen in Figure 4.7.

Power system

The functions of the power system also known as the Energy Power System(EPS) is to generate, store, regulate and distribute electrical power [12]. Itconsists of three main elements: a power source, an energy source, and a powerdistribution, regulation and control unit system.

According to Poghosyan and Golkar [15] photovoltaic solar cells are theprimary energy source for most CubeSats utilizing state-of-the-art triple-junctionsolar cells achieving efficiencies between 27-33%. Taking into considerationthat this is the most used power source among CubeSat developers, havingmultiple options to choose from increases competition which tends to lowerthe prices, has lead to the author choosing it as a power source for the designof this CubeSat.

Furthermore Poghosyan and Golkar [15] argues that given the CubeSatconstraints on surface area the solar cells can be either body mounted or de-ployable, the later increases the risk of mission failure as there is a possibilityof them not deploying in-orbit. On the other hand, it does come with advan-tages such as protecting the solar cells inside the structure at launch and itincreases the power generation, easily reaching up to 30 W.


Figure 4.7: The CubeSat specification in the framework of small satellite clas-sifications. As can be seen CubeSats are classified based on mass and volume,figure was retrieved from [15] .

Considering energy power storage, it is usually needed because of the con-straints given by solar eclipses and peak power requirement of the payload(s).Several possibilities exist as for example high energy density lithium ion bat-teries and lithium polymer batteries [15].

The last element (power distribution, regulation and control unit system) isnormally custom built by the satellite providers [15] in relation to their systemrequirements.

In conclusion, today it is common practise to buy the power system as awhole, it is then refereed to as the EPS, by satellite providers such as BCT,ÅAC Clyde Space and GomSpace [15], overall they can deliver EPS subsys-tems for CubeSats with adequate capabilities based on solar power and battery


storage and therefore the author has chosen to integrate a full EPS system in-stead of choosing each element separately.

Guidance, navigation and control

The GNC, also known as the Attitude and Orbit Control subsystem (AOCS), ismade up of two main elements: the Orbit Determination and Control System(ODCS) and the Attitude Determination and Control System (ADCS) [15].

As Poghosyan and Golkar [15] continue to explain, these two subsystemsare equally important. The ODCS measures and maintains the position of theCubeSat’s centre of mass as a function of time, as the CubeSat will be in LEO aglobal navigation satellite system (GNSS) will provide position determination.The ADCS measures and maintains the satellite’s orientation about its centreof mass. It does this by using sensors such as magnetometers and sun sensorsto determine the CubeSat’s attitude and actuators such as magnetorquers andreaction wheels to stabilize and orient the CubeSat in the right direction [15].

As was the fact for the EPS system, also the GNC is offered as an integratedunit providing 3-axis control for CubeSats. In general, the CubeSat GNC is amature technology with many flight proven highly integrated systems availableon the market [15].

Operational modes which the GNC subsystem is envisioned to be able todeliver are the following:

• Nadir pointing: the spacecraft is controlled w.r.t to the inertial referenceframe for nadir pointing of ground track target this is for down- anduplink of TT & C and most importantly for receiving and transmittingVDES data by the payload.

• Sun pointing: Spacecraft is controlled w.r.t the inertial reference frameto point the solar panels at the sun for maximum solar input.

• Inter satellite pointing: point to satellite in track or adjacent orbital planefor ISL communication.

• De-tumbling mode: reduces spacecraft’s angular rates at separation fromlaunch vehicle as well as functions as a fall-back safe mode during anoma-lies.

• Non-operation mode: satellite’s attitude is uncontrolled, to for examplesave battery.


Thermal control system

The subsystem thermal control keeps all components and payload(s) withinoperating temperature ranges during normal operations and keeps everythingwithin survival limit under all circumstances [12]. It does this by balancingthe heat input in relation to the energy that is emitted from the satellite. Thisis of importance since a satellite in LEO orbit experiences in short amount oftime extreme temperature fluctuations from -100 degrees Celsius in eclipse to+100 degrees Celsius in direct sunlight [15].

The thermal control system of a CubeSat has generally been composedby a passive thermal control system, which is defined by having no powerinput. Examples of passive thermal control systems are Multi Layer Insulation(MLI), thermal coating, sun shields, thermal straps, louvers, radiators and heatpipes [15]. A passive thermal control system is usually reliable, has low mass,volume and cost [15]. Hence it is a preferable option over an active thermalcontrol systems, when considering that CubeSats have restricted power, massand volume. The author has for that reason chosen to proceed with a passivethermal control system.

Communications

The communications subsystem works as the interface between the ground-and space segment, the space segment and the subject (the vessel segment inthe VDES system) and between the satellites (ISL) in the space segment. Itcould possibly also function as the interface between the space segment and thebase stations already employed by Saab (shore segment in the VDES system),see section 2.1 and Figure 4.2 where design options have been evaluated.

The communications subsystem is composed of receivers, transmitters andantennas [12].

Several links are proposed for the system:

• Feed link

• User link

• ISL

The feed link acts as the interface between the ground segment and thespace segment (ground station - satellite, satellite - ground station) providingTT & C as well as payload data. According to Poghosyan and Golkar [15]the capability of communication is increasing for CubeSats from a few Mbpsthrough S-band today, which could be a limiting factor, to higher data rates


using the higher X-band of several hundred Mbps. As state-of-the-art is al-ready being employed for technology demonstrations the CubeSat standardcould rather soon be capable of such high-band throughput.

The user link is the interface between the space segment and the subject(satellite - ship, ship - satellite), using VHF radio on VDE channels for up anddown-link as well as the ASM and AIS channels for up-link from subject tospace segment.

The ISL (satellite - satellite) will function as a link layer of the space seg-ment to increase the speed of downlink to the subject and the ground segmentby using routing to establish which satellite that has a direct link to the targetin order to transfer through the link layer to that specific satellite, of which thetarget is in the FOV, for downlink of data.

An example of ISL can be seen in Figure 3.2 where satellite 1 transfersdata to satellite 2 in order to downlink vessel data to the ground segment tobe sent to the user at shore by IP, it could of course be the other way around,such that the data is sent from user at shore by IP to be uploaded by groundsegment to the space segment and sent through the space segment by ISL tobe downloaded to the targeted vessel.

Command and data handling

The C & DH subsystem is responsible for receiving, validating, decoding anddistributing commands to other subsystems. Moreover, it gathers prepares andstores house keeping and mission data for downlink or on-board utilization[15].

Commonly used on-board data handling systems among CubeSats devel-opers are micro-controllers based on high performance and power efficientARM processors, however, sensitive to space radiation, and field programmablegate arrays (FPGAs) with space heritage and easily integrated with on-chipmemories [16].

CubeSat data storage requires high reliability and a variety of differentmemory technologies have been developed, but maybe most interesting is theflash memory technology which can store up to hundreds of gigabytes [16].

In conclusion, the fundamental limiting factor for CubeSats, again accord-ing to Poghosyan and Golkar [15], is not the on-board storage, but the datadownlink capability. C & DH is a mature field with several available options,nevertheless radiation hardened components will be required to extend the life-time of CubeSats in LEO to longer than a year [16].


4.3.4 Choice of satellite providerWhen choosing the satellite provider, the payload physical requirements wereaddressed as discussed earlier in Table 4.4. Which provider could fulfil thepayload requirements through the smallest CubeSat possible? It had alreadybeen decided that a 1 U CubeSat would not be enough due to size constraintsof the payload. Also the Technology Readiness Level (TRL) was addressed,where only a TRL of 9, which means that the technology has previously flownin space, was accepted to be further investigated, see Appendix E in report”State of the Art: Small Spacecraft Technology” NASA [16] for further de-scription on TRL.

In this analysis the satellite providers listed in NASA [16] as CubeSatproviders have been considered, as can be seen in Table 4.5. At several pointsthe actual data has been checked and at times altered to be correct accordingto the respective satellite provider’s web page. Moreover, only the satellitesthat clearly met the three physical payload requirements of mass, dimensionand power were chosen for further studies. The conclusion from this analysis,was as can be seen in the table, to continue the analysis with the 6 U bus fromÅAC Clyde Space and 6 U bus from ISIS. Nevertheless, BCT, GomSpaceand NanoAvionics (NA) could be an option, but due to lack of informationregarding payload power capability, an important aspect since as of writingthe payload power requirement is 20 W OAP and 60 W PP, see again Table4.4. Taking this into consideration the choice was made to not study theseplatforms any further.

There are reasons for which one may argue to only choose the ISIS 6 Uplatform since it shows the capability to carry the mass of the 1 kg payloadwhereas ÅAC Clyde Space’s 6 U bus does not, however, the author assumesafter having seen all the other platforms payload mass capabilities that alsoÅAC Clyde Space’s buses should have this capability. Moreover, ÅAC ClydeSpace also has its 3 U spacecraft bus as an option if the payload power canbe optimized just a bit further. Here a decision has to be made whether thedevelopment of the satellite platform should aim for several payloads and thusutilize the 6 U platform or only use the single VDES payload and essentiallyallow a 3 U CubeSat to be used. Both alternatives might have their advantagesand disadvantages considering costs, operations, availability, etcetera.

Having taken all of this into account, the further investigation of key bud-gets was chosen to be done with the 6 U platform provided by ISIS Space.Thus leaving out ÅAC Clyde Space’s promising 6 U satellite platform, mainlybecause of lack of open source data of their products to have been able to


construct the design budgets, see Table 6.6 for the constructed design budget.Furthermore, only the VDES payload will be used allowing for full control andavailability of satellites for VDES purposes, as the system requirements, seeTable A.2 and requirement FR-9 requires the system to have low data latencyand to minimize the number of used satellites in the constellation.

Table 4.5: Potential satellite providers for CubeSats collected from [16] forcomparison with the payload requirements from Table 4.4 that drives the de-sign of the satellite.

Vendor Satellitedimension

Payloadmass [kg]

Payloaddimension [U]

Payloadpower [W] TRL Add

Ref Comment

ÅAC Clyde Space 3 U - 1.5 U 18 OAP, 60 PP 9 [60] Nearlymeets REQ.

6 U - 4 U 60 OAP, 100 PP 9 [60] Meets REQ.

AMA 3 U 4 - 12 OAP - - MissingTRL level.

6 U 12 - 20 OAP - - MissingTRL level.

BCT 3 U 1.5 1.5 U 60 OAP - total ,6.3 OAP - bus 9 [61] Low power.

6 U 1.5 4 U 140 OAP - total,6.3 OAP - bus 9 [62] Possibly, power

might be too low.

GomSpace 3 U - 1 U 8-15 OAP 9 [63] Low Power.

6 U - 2 U 12-24 OAP 9 [63] Possibly, powermight be too low.

ISIS3 U(advanced bus) 1-2 1,5 U 5 OAP 9 [64] Low power.

6 U(advanced bus) 5.5 2.5 U 20 OAP 9 [65] Meets REQ.

MSS 6 U - - 16 OAP - total 9 [66] Lack data,low power.

NA 3 U 4 Up to 2 U 75 OAP - total 9 [67] Possibly, powermight be too low.

6 U 7.5 Up to 4 U 75 OAP - total 8 [68] Possibly, still notproven in flight.

Pumpkin 3 U 2 2 U - 9 - Not enough data.

TNSS 6 U 3 3 U 180 PP 9 [69] Does not providedata regarding OAP.

UTIAS SFL3 U 1 - 2 9 - Not enough

payload power.

8 U 2 - 4 9 - Not enoughpayload power.

16 U 6 - 45 9 - Enough power,but 16 U.

Note: Additional Reference (Add Ref), Requirement (REQ), total - power for whole satellite, bus - power for

satellite bus.


4.4 Ground SegmentA ground segment consists of a network of ground stations and control centres.It supports the space segment, relaying the mission data to the users of thesystem. It also collects telemetry data beneficial to command and control thesatellite bus and payload. Examples of such is monitor of spacecraft health,track of its position and report of attitude [16].

4.4.1 Ground station networkInitially to explain the features of the ground station network which is the thereceiving and transmitting system on ground in relation to the space segment,in this case the satellite constellation. The ground station is composed of fivemain components: a controllable antenna for receiving telemetry and payloaddata from the space segment and to transmit commands, an antenna controlsystem in order to track the passing satellite, an antenna feed, a receiver unitand a transmit unit [70].

The ground stations are connected to each other by internet offering capa-bilities to rapidly send the payload data to end users or to send and receive TT& C data from and to the MCC where the satellite operations are performed[70]. It is also possible to quickly send data to the data center for enhance-ment of the data in order to for example provide a web interface where theusers/customers can track vessels and gain information of e.g identity, posi-tion, velocity, route, fuel consumption, etcetera.

4.4.2 Satellite operationsFor small spacecraft operations several different operations centres such asSatellite Operations Control Center (SOCC) that normally controls the satel-lites, Payload Operations Control Center (POCC) controlling the payload(s)and MCC focusing on planning and operating the mission are usually the same[16].

In a similar manner one type of operations centre is used for the satelliteVDES system and has been refereed to as MCC in this work , see Section 4.1.


4.5 Launch segment

4.5.1 Launch vehicleHere follows a short explanation to the launch vehicle, possible launch providersand emerging technologies within this field, enhancing the possibility for smallsatellite missions.

According to NASA [16] of the 464 spacecraft launched in 2017, 62 % ofthese were nano satellites or lesser, that corresponds to 289 spacecraft. Thistrend has been forecast to continue the upcoming years causing a major impacton launch vehicles, integration and deployment systems to adapt and developtheir capability in order to meet the advancement in missions and technologywhich is on the horizon for nano satellites.

There are three categories of interest for launching small spacecraft intoorbit [16].

• Ride-share as secondary payload

• Dedicated ride-share

• Dedicated launch vehicle

Ride-share as secondary payload, also known as piggybacking, means thatthe launch vehicle customer (primary payload) decides whether it is possibleto share a ride and if so when the secondary payload can be dispensed. It isalso decided by the primary payload at what altitude and inclination the launchvehicle will eject its payload [16]. This is perhaps the most common way oflaunching a small satellite into orbit.

Furthermore, identifying a secondary payload launch is done by a launchintegration company, these companies purchase spare space on a launch vehi-cle and try to integrate as many small spacecraft as possible. More often thennot they are the same as the satellite providers, see subsection 3.2.1, companiessuch as Tyvak Nano-Satellite Systems (TNSS), Surrey Satellite TechnologyLTD, UTIAS SFL and ISIS, to mention a few [16].

Dedicated ride-share missions are an upcoming trend defined by having athird party integrator purchase an entire launch from a launch vehicle provider.It then, in the absence of a primary payload, is free to integrate and contractmultiple small satellites as seen suitable for the specific mission [16].

One of the major examples of this was the SSO-A Smallsat Express mis-sion synchronized by the third party Spaceflight Services, it was launched inDecember 2018 on a falcon 9 rocket. It is to date the largest single ride-share


mission from a United States based launch vehicle, deploying 64 spacecraftfrom 34 different organizations, whereas 49 were CubeSats, to a SSO in LEO[71].

Nevertheless, the main disadvantage of launching as a secondary payloadin a ride-share mission or even as one of many primary payloads on a ded-icated ride-share mission is the inability to launch into your specific desiredorbit. One way to counter this is to use an orbital manoeuvring system, such asthe Shuttle Expendable Rocket for Payload Augmentation (SHERPA), to forexample perform LEO altitude shifts or change of inclination [16]. SHERPAperformed its maiden flight on the mentioned SSO-A Smallsat Express mis-sion [71].

Lastly, and perhaps in the future of most interest is the alternative of adedicated launch which enables launching into an orbit of chosen altitude, in-clination and eccentricity for small spacecraft. At the time of writing a marketof small launchers is not well established, however, there are a few alternativesthat have reached a TRL of 9 such as Rocket Lab with its Electron launcherand Northrup Grumman Innovation Systems with its Pegasus, Minatour 1 andMinatour 5 launch systems [16].

Rocket Lab offers launches with its newly developed launch vehicle Elec-tron designed to lift up to 150 kg to 500 km altitude in a SSO, but also withthe capability to tailor circular or elliptical orbits between an inclination of 45and 98. Its second launch occurred in 2018 successfully placing 6 CubeSatsin LEO [16].

Investigating further into a launch with Electron, it has a very exact or-bit injection accuracy, although the main concern for the mission is that thesatellite(s) survive the acceleration loads and random vibrations at launch. Re-ferring to Beck [72] the max lateral loads lay in-between+/−2 g and the axial+8/− 4 g, thus rigorous vibrational testing is needed.


4.5.2 Separation systemAnother aspect of the CubeSat standardization is the dispenser, a deploymentsystem, which earlier has been noted as the separation system see Figure 4.2.It functions as the interface between the CubeSat and the launch vehicle. Thisit does by attaching the satellite to the launch vehicle, reducing the mechanicalloads on the satellite during launch and by releasing the satellite into space. Itsprimary mission, nevertheless, is to ensure safety to the CubeSat and protectthe launch vehicle, main payload and other CubeSats during launch [73].

Dispensers have been standardized to align with the standard CubeSat formfactor and are provided by several companies depending on the size of theCubeSat and depending on what specific launch vehicle is aimed to be used[58]. A few examples of such providers have been listed earlier, see subsection3.2.1.

To give an idea of the working mechanism, a deployment signal is sentfrom the launch vehicle to initiate the release mechanism to open the door andto eject the CubeSat by utilising a spring mechanism [58].

An example of several 6 U dispensers can be seen in Figure 4.8.

Figure 4.8: Three 6 U Dispensers (from left to right) provided by PlanetarySystems corporation, ISIS and TNSS, figure was retrieved from [73].

Chapter 5

Market analysis

5.1 Current AIS constellationsAt the time of writing already many companies are operating or planning tolaunch a constellation for Satellite AIS, the ones listed in Table 5.1 have beenidentified through own research or from the database provided by Kulu [6], [7]and [74] in particular, where all satellites of non importance have been filteredby listing AIS as filter option for the column called Constellation field.

Occasionally there have been contradicting information, mostly becauseKulu [74] lists all launched satellites by respective service provider includingde-orbited ones and not only the actual active satellites, which is what hasbeen pursued to be showcased here. Hence other sources from own researchhas paid a significant part to the result, as can be seen in Table 5.1 whereadditional sources are listed in the column ”Reference”.

Worth mentioning, is that currently only UK Space (VESTA mission), byorder from ExactEarth, and Space Norway (NorSat-2) have operative satellitespublicly stating to be performing tests for VDES. All the other satellites andfuture satellites listed in Table 5.1 are/will be using AIS, which as explainedearlier in Section 2.1 is limited to receive only.

Moreover, constellations of greater value to their customers, having launchedenough satellites for acceptable data latency are Spire Global, ExactEarth andORBCOMM, where Spire Global is the company with the most operativesatellites for AIS.

Another fact, looking again at Table 5.1 that attracts one’s attention is thatGomSpace has been very active on this market both as a satellite provider aswell as payload provider, having signed contracts for several major future con-stellations, e.g Sky and Space (SAS) Global and AISTech to provide several

64

CHAPTER 5. MARKET ANALYSIS 65

Table 5.1: Shows service companies providing satellite AIS constellations.

ServiceProvider

SAT-AISin orbit

SAT-AIS plannedin orbit Satellite provider Payload provider Reference

Space Norway 4 N/A UTIAS SFL Kongsberg Maritime [26]

ExactEarth 67 67UTIAS SFLand Thales

Alenia Space

(Kongsberg Maritime andSRT Marine) and Harris [6], [75]

UK Space 1 N/A Surrey SatelliteTechnology Ltd Honeywell [76]

Spire Global 76 175 Spire Global Spire Global [77],[78]

ORBCOMM 18 N/A Sierra Nevadaand ÅAC Clyde Space

Boeing Spaceand ÅAC Clyde Space [79], [80], [81]

Hawkeye 360(Pathfinder) 3 30 UTIAS SFL GomSpace [82],[83]

Sky and Space Global 3 200 GomSpace GomSpace [84], [85]AISTech 2 150 GomSpace GomSpace [86], [7]

Commsat Technology 7 72 Commsat Technology Commsat Technology [87]Aerial Maritime 0 104 GomSpace GomSpace [88]

Kleos Space 0 20 GomSpace GomSpace [89], [90]SRT Marine 0 6 ÅAC Clyde Space SRT Marine [91], [92]

Tekever 0 12 Tekever GMV [93]Karten Space 0 14 Karten Space N/A [6]

Total 181 838Note: Not Announced (N/A)

hundred satellites.In the two sections earlier Space Norway and their VDES attempts have

been explained as well as ExactEarth and their constellation, see Section 2.6.1- 2.6.2.

5.2 AIS CubeSats launchedTo further explain what kind of satellites are being launched for AIS/VDESpurposes, it is showcased in Table 5.2 several key aspects of such. They havebeen listed since they are provided by service providers from Table 5.1 forAIS/VDES operations and in CubeSat form factor. Thus they are of specialinterest to this work, showcasing what limitations and possibilities exist andwhat is common practise in the design of such satellites. This has been takeninto account when developing the design budgets for a typical satellite in thesatellite VDES system, as can be seen in Table 6.6.

66 CHAPTER 5. MARKET ANALYSIS

Table 5.2: Shows satellites that have been launched providing AIS/VDES ca-pabilities.

Name of CubeSat NorSat - 2 Lemur 2 Vesta Pathfinder,Hawk Diamond EV-9 AISTechSat

Company/Agency,Country

Space Norway,Norway

Spire Global,US

UK Space,UK

Hawkeye360,US

SAS Global,Israel

ExactEarth,Canada

AISTECH,Spain

Mission Tech Demo Commercial Tech Demo Tech Demo Commercial Commercial CommercialForm factor 24 U 3 U 3 U 20 U 3 U 8 U 2 U

Payload(s)AISVDESCamera

AISADS-B,GNSS-RO

VDES SDRRF front end

AISADS-BM2M/IoT

AISAISADS-BIoT/M2M

ISL No No N/A Yes Yes, 4-ways No N/AOrbitalaltitude 600 km 400-650 km 575 km 550-650 km 500 km 650 km 500-600 km

Inclination SSO98.7 deg

SSOEO

SSO97.77 deg SSO EO

0 +/- 20 degEO6 deg SSO

Mass 15.6 kg 4 kg 4 kg 15 kg 10 kg 5.5 kg N/A

Communications S-Band/UHF UHF/S-band S-band S-band/X-band S-band S-band N/A

Launchvehicle

Piggyback,Soyuz

Piggyback,PSLV, ISS

Ride-share,Falcon 9

Ride-share,Falcon 9

Ride-share,PSLV

Piggyback,PSLV

Ride-share,Falcon 9

Year of(first) launch 2017 2015 2018 2018 2017 2015 2018

Deploymentsystem XPOD Duo N/A N/A XPOD N/A XPOD N/A

Ground stations 2 dedicated 30+ dedicated N/A K-SAT N/A 1 dedicated N/ASelf sustained No No No Yes No No NoSatellite lifetime 5+ years 2 years N/A N/A 4 years N/A N/A

Pointing control 3 - axis+/- 0.5 deg N/A 3 - axis 3 - axis 3 - axis 3 - axis N/A

Power source Solar power48 W OAP Solar power Solar power Solar power Solar power Solar power

6-14 W OAP Solar power

Reference [8] [47], [94], [95] [76] [82] [85], [84], [96] [9] [97],[98]Note: Not Announced (N/A), Technological Demonstration (Tech Demo).

Chapter 6

Results

6.1 Stakeholder analysisThe key stakeholders were, as earlier mentioned, identified as the customers,the payload provider, the satellite provider, the service provider and the satel-lite operator.

From the key stakeholders the stakeholder requirements could be imple-mented. A stakeholder requirement can be explained as a statement for a stake-holder need. It is a problem-oriented statement [5].

In this case the problem is specified by overloaded AIS channels and inca-pability of tracking and communicating with vessels at open sea.

6.2 Technical requirementsThis section explains the technical requirements, starting with stakeholder re-quirements also refereed to as user requirements.

6.2.1 Stakeholder requirementsThe stakeholder requirements have been outlined by having the key stakehold-ers in mind. The produced requirements can be seen in Table 6.1.

6.2.2 Mission requirementsFollowing the stakeholder requirements that represent the stakeholder needs,the mission requirements could be outlined for the system, see Table 6.2. The

67

68 CHAPTER 6. RESULTS

Table 6.1: Stakeholder requirements in the satellite VDES system, whereeach requirement has been defined either as characteristic (Char) or capability(Cap).

Identifier Stakeholder Statement Cap/Char

SH-1.1 Serviceprovider

The system shall allow fortracking of vessels carryingVDES/AIS transponders globally.

Char.

SH-1.2 Serviceprovider

The system shall providecontinuous global coverage. Cap.

SH-2.1 CustomerThe system shall not bemore expensive than theAIS system in use today.

Char.

SH-2.2 Customer The system shall be secureand not easily jammed. Char.

SH-3.1 Payloadprovider

The system shall allow fordirect communication between vessels,satellites and in-between.

Cap.


The SDR-platform shall havecompatibility for satellitesas well as UAVs.

Cap.


The system shallhave a down - and uplinkin the maritime frequency domain.

Cap.


The SDR-platform needed forthe system shall be developedwith limited investments from the payload provider.

Char.


The satellite platform shall bedeveloped in short timeand to as low cost as possible.

Char.

SH-4.0 Satelliteoperator

The satellite constellation shall functionautonomously in nominal operations. Cap.

SH-5.0 Satelliteprovider

The SDR platform shall notexceed the dimensions, mass constraints andpower constraints specified by the satellite provider.

Char.

CHAPTER 6. RESULTS 69

mission requirements represent the acceptance criteria of the mission, if theseare fulfilled the mission can be agreed to have been a success [5].

Table 6.2: Mission requirements in the satellite VDES system influenced bywhat has been done by Øystein et al. [10].

REQ ID Description Statement

MR-01 Geographic coverage

The system shall providecontinuous global coveragefor communication and tracking with/ofvessels carrying VDES/AIS transponders.

MR-02 Timeliness The system shall have aone-way latency delay of <2 s.

MR-02 Validation

The system shall utilizeavailable information for validationof positional informationin the received VDES messages.

MR-03 Data storage The system shall allow foron board storage and processing of data.

MR-04 Flexibility

The system shall allow futurechanges to the VDES system,e.g changes infrequency and/or signal format.

MR-05 CostThe system shall not be moreexpensive for the user than theAIS system in use today.

MR-06 Security The system shall not bejammable by conventional means.


6.2.3 System requirementsThere are a few differences between system- and stakeholder requirements ac-cording to Douglass [99], the main difference is that system requirements aremore focused on:

• System properties

• Engineering language

• To be quantitative

The system requirements have been defined using the approach described byNASA [11], whereas five main categories have been established: system In-terface Requirements (IR), system Constraint Requirements (CR), Functionalsystem Requirements (FR), system Performance Requirements (PR), systemEnvironment Requirements (ER), see Table A.1, A.2, A.3, A.4 and A.5 in theAppendix.

6.3 Constellation analysisHere follows a short analysis on the envisioned satellite constellation. Theconstellation considered is a polar or near polar constellation of common al-titude and spacing between orbital planes, except for the non-synchronizedorbital planes, see again Figure 4.5.

To begin with, the altitude and minimum elevation angle is changed inorder to get an idea of how the swath width is affected, this can be seen inFigure 6.1. The minimum elevation angle is of course dependent on the pos-sible antenna gain of the satellite’s antenna(s), which is power dependent, aswell as the antennas of the ground stations. Moreover, the results of Figure6.1 corresponds well with what has been done in George et al. [57].

Considering an orbital altitude of H = 550 km, in order to gain mostpossible coverage, as the coverage increases with altitude as can be seen inFigure 6.1, to avoid most of the aerodynamic drag to stay in orbit 5+ yearswithout any propulsion system and to still be able to de-orbit within requiredtime of 25 years after end of operational life.

Why be concerned about the de-orbit time? Well, according to Qiao, Ri-zos, and G. Dempster [100] there is a chance of not de-orbiting within theframe of 25 years after end-of life in an orbit higher than 550 km for a 6 UCubeSat of a corresponding mass of 6 kg, with a drag coefficient of 2.2 andin a SSO. Their analysis was done using the STK 9.0 lifetime tool and can


400 450 500 550 600 650 700 750 800

Altitude [km]

2500

3000

3500

4000

4500

5000

5500

6000

6500

Sw

ath

wid

th [k

m]

Altitude vs. swath width, when altering the minimum elevation angle

elevation angle 0°

elevation angle 1°

elevation angle 2°

elevation angle 3°

elevation angle 4°

elevation angle 5°

elevation angle 6°

elevation angle 7°

elevation angle 8°

elevation angle 9°

elevation angle 10°

Figure 6.1: Swath width as a function of orbit altitude for minimum elevationangle 0-10 deg, the calculations are shown in Appendix D.

be seen in Figure 6.2 and was of interest to this study since it complies to aCubeSat design very similar to our own, see subsection 6.4.1.

Furthermore, considering an elevation of ϵ = 0 the number of satellitesfor the considered altitudes could be estimated as can be seen in in Table 6.3.Such that for an altitude of 550 km approximately 32 satellites are needed forfull global coverage, where at least one satellite is always seen by an observeron the surface of Earth. To provide further details on the constellation, thecalculations gave 8 satellites trailing in each orbital plane, with a total of 4orbital planes. The approach that has been used for the calculations can beseen in Section 2.4.2, equation 2.5 - equation 2.15.

Looking instead at Table 6.4, many more satellites tend to be needed forthe same altitudes when the minimum elevation angle is increased. This isworth paying attention to since many antennas are limited to around 5 - 10deg of elevation minimum angle. For a minimum elevation angle of 10 degreesaround 91 satellites are needed at an altitude of 550 km for global continuouscoverage.


Figure 6.2: Orbital lifetime as a function of initial orbit altitude and CubeSatform factor. Each satellite has the same mass as its respective form factor,figure was retrieved from [100].

Table 6.3: Showcases how the minimum number of satellites Ntot needed forglobal coverage changes with orbital altitude H and which velocities vc, pe-riods T and slant ranges D0 are to be expected when keeping the minimumelevation angle at zero, the calculations are shown in Appendix D.

H [km] Ntot [-] vc [km/s] T [min] D0 [km]400 50 7.67 92.4 2294.0450 45 7.64 93.4 2437.8500 45 7.62 94.5 2574.5550 32 7.58 95.5 2705.2600 32 7.56 96.5 2830.8650 32 7.53 97.6 2951.9700 32 7.51 98.6 3069.0


Table 6.4: The number of satellitesNtot needed for continuous global coverageand how it corresponds to the number of orbital planesNp and satellites in eachplane Ns. The minimum elevation angle is kept at 10 degrees. The slant rangedecreases with increasing elevation angle, see again Appendix D.

H [km] Ntot[-] Ns [-] Np [-] D [km]400 120 15 8 1439.8425 120 15 8 1505.6450 98 14 7 1570.0475 98 14 7 1633.2500 91 13 7 1695.1525 91 13 7 1755.9550 91 13 7 1815.6575 72 12 6 1874.4600 72 12 6 1932.2625 72 12 6 1989.2650 66 11 6 2045.3675 66 11 6 2100.7700 66 11 6 2155.3

6.3.1 Inclination of a sun synchronous orbitIn the case of launching an experimental satellite due to power requirements,as earlier discussed in Section 2.4.6, a SSO giving full-time illumination couldbe desired.

As can be seen by equation 2.24 the inclination required to have a SSO canbe calculated for different altitudes in circular LEO, Table 6.5 showcases theresult.


Table 6.5: This table gives an indication of what inclinations are needed toestablish a sun synchronous orbit for the altitudes in question.

H [km] i[]

400 97.03450 97.22500 97.40550 97.60600 97.79650 97.99700 98.19

6.4 Design results

6.4.1 Design budgetHaving considered Section 4.3.2-4.3.4, the design budget was developed inorder to fit a 6 Unit structure CubeSat provided by ISIS. The result was a totalmass of 5.8 kg, a dimension of all the systems with a total of 3.7 dm3 and atotal OAP of 70 W as can be seen in Table 6.6.

Note that three On Board Computers (OBCs) are used, to designate eachone with a specific task: the House Keeping Computer (HKC) taking care ofanything that has to do with analysing telemetry of the satellite, the AttitudeDetermination and Control Computer (ADCC) to take care of pointing and at-titude calculations and lastly the Payload On Board Computer (POBC) to dealwith all on board processing regarding payload operations from the satellite,such as processing of VDES data, decoding messages, designating a time slotin the MCS, etcetera.


Table 6.6: Table showcases the design budget for a VDES satellite using ISIS6 U spacecraft bus, pay attention the total mass, volume and power.

Element Specific Product Mass [kg] Volume [dm^3] Power [W] Ref

Payload SDR platform Saab - adaptedversion of R5A 1 1.4406 20 [48]

AOCSADCS(3-axis magnetorquer+ 3-axis magnetometer)

iMTQ 0.196 0.14689003 1.2 [101]

3-axis reaction wheel NanoTorque GSW-600 0.94 0.550525 2.55 [102]

2 Fine sun sensors Nano sense FSS 0.004 0.00242 0.0268 [103]

GNSS receiver OEM719 0.031 0.036432 1.32 [104]

Communications Feed link transmitter ISIS High data rateS-band transmitter 0.3 0.28512 9.2 [105]

Feed link receiver ISIS UHF/VHFtransceiver TRXUV 0.085 0.1296 0.2 [105]

ISL transceiver NanoCom SR2000 0.271 0.21342 5.94 [106]

GNSS antenna Tallysman TW1322 0.026 0.00042 0.1 [107]

ISL Antenna NanoCom ANT2000 0.11 0.1930404 10.7 [108]

Payload antenna Yagi-Uda antenna 8 dBi 0.3 6.2 × 6.2 × 7.3 1 [8]

Antenna system ISIS deployableantenna system 0.085 0.067228 0.24 [109]

C & DH 3 OBC ISIS OBC 0.2 0.214272 1.2 [110]

Thermal control Passive thermal control MLI - - - -

Power system Deployable solar panels CubeSat solar panels 0.3 - - [111]

Li-Ion battery pack NanoPower BPX 0.5 0.327918 6 [112]

Power distribution Nanopower_P60_acu 0.054 0.00321276 1.695 [113]

Nanopower_P60_pcdu 0.057 0.012363632 0.165 [114]

Nanopower_P60_dock 0.08 0.00864 5.35 [115]

Structures andMechanisms Spacecraft Bus 6 U structure 1.1 7.705515 - [116]

Margin add an extra 5% 0.28195 0.181605091 3.34434 [12]

Total 5.9 3.8 70.2


6.4.2 Link and data budgetsHere follows the results of a link- and data budget between the space segmentand the user segment. Moreover, a few communication parameters are listedas they have been used in the link- and data budget, see Table 6.7.

In regard of the link budget: the satellite links to a single ship and thequality of the linking for both up- and downlink are estimated, see Tables B.1- B.2 in the Appendix and Figure 6.3.

Later on, the antenna gains of the vessel and satellite change with the el-evation angle as can be seen in Table 6.8, as indicated by ITU-R [117]. Notethat the satellite antenna main beam is pointed towards the horizon at an offset-nadir angle of around 66-67 degrees, thus the highest antenna gain is at ϵ = 0

degrees and then decreases with increasing elevation angles. This gives thefollowing results for the link budgets, see Tables B.3- B.4 in the Appendixand Figure 6.4. Here the energy per bit to noise spectral density ratio, Eb

N0, is

not taken into account since several modulation schemes are considered byITU [117] with different data rates. They have also provided the thresholdsfor the carrier to noise density ratios, C

N0, shown in the figure for the different

modulation schemes in order to compare the performance of the linking [117].Lastly, concerning the data budget, the time of contact,Tc, for a satellite

passing directly over a single vessel and the data amount, DA, retrieved fromit are plotted against the minimum elevation angle, see Figure 6.5. In relationto Figure 6.5, a few results are highlighted in a data budget, see Table B.5 inthe Appendix.

Table 6.7: Satellite communications parameters for SAT-VDE downlink anduplink between subject and satellite used in link- and data budget.

Constants Downlink Uplink ReferencesCarrier frequency [MHz] 162 157 [19]Pt [W] 1 6 [8],[117]Gt [dB] 8 3 [117]Gr [dB] 3 8 [117]LP [dB] 3 3 [117]Ts [dBK] 30.2 25.7 [117]BR [kb/s] 240 240 [20]


0 10 20 30 40 50 60 70 80 90

Elevation angle [°]

60

70

80

90C

/N0 [d

BH

z]

Carrier to noise density ratio vs. elvation angle

Downlink: C/N0

Uplink: C/N0

0 10 20 30 40 50 60 70 80 90


0

10

20

30

40

Eb/

N0 [d

B]

Energy per bit to noise spectral density ratio vs. elvation angle

Downlink: Eb/N

0

Uplink: Eb/N

0

Figure 6.3: Two different performance indicators, CN0

and Eb

N0, for the satel-

lite communication system plotted against the elevation angle. The resultsare shown in Tables B.1 - B.2 in Appendix B and calculations are shown inAppendix D.

Table 6.8: The satellite antenna gain and the ship antenna in relation to theelevation angle, as according to ITU-R [117].

ϵ [] Gsat [dB] Gship [dB]0 8 310 8 320 8 2.530 7.8 140 6.9 050 5.5 -1.560 3.6 -370 0.7 -480 -2.2 -1090 -5.5 -20


0 10 20 30 40 50 60 70 80 90


30

40

50

60

70

80

90

C/N

0 [d

BH

z]

Uplink: carrier to noise density ratio vs. elvation angle

C/N0

QPSK/CDMA/4 QPSK

8PSK16QAM

0 10 20 30 40 50 60 70 80 90


30

35

40

45

50

55

60

65

70

C/N

0 [d

BH

z]

Downlink: carrier to noise density ratio vs. elvation angle

C/N0

BPSK/CDMA/4 QPSK

8PSK

Figure 6.4: When the antenna gains are not kept constant the carrier to noisedensity ratio, C

N0, can be seen increasing at first to an elevation of 40 degrees

and then to decrease. Results are shown in Tables B.3 - B.4 in Appendix Band calculations are shown in Appendix D.


0 10 20 30 40 50 60 70 80 90

Minimum elevation angle [°]

0

5

10

15

Tim

e of

pas

sage

[min

] Time of contact vs. minimum elevation angle

0 10 20 30 40 50 60 70 80 90

Minimum elevation angle [°]

0

50

100

150

200

data

am

mou

nt [M

b]

Data ammount vs. minimum elevation angle

Figure 6.5: The time of satellite contact, Tc, and the data amount, DA, col-lected plotted versus the minimum elevation angle, ϵ, in regard to a singlevessel. The calculations can be seen in Appendix D and a Table of the resultsis given in Table B.5, Appendix B.


6.4.3 Mission designThe preliminary mission design for a Satellite VDES can be seen in Table 6.9, have in mind that the mission proposal can change as the system evolves.

Table 6.9: The preliminary mission design is showcased here.

Type Preliminary mission design

Mission Satellite VDESOrbital altitude Common altitude: 550 kmInclination Common inclination: polar - 90 degMinimumelevation angle 10 deg

Number of satellites 91 satellites needed for continuous global coverageSatellitelifetime 3-5 years

Satellite mass 5.9 kgPower consumption 70.2 W (OAP)

Spacecraft dimension 6 U (100 × 226.3 × 340.5 mm^3) - platform,3.8 U used for payload and all subsystems

Payload SDR-platform for VDESCommunication Feed link: S-Band, ISL: S-Band, user link: VHF

On board computerHKC - for housekeeping tasks,ACC - for pointing and controlling attitude,POBC - for payload operations

Self sustained NoAttitude 3-axis stabilizedPointing speed 3 deg/s

Power system Power generation: deployable solar panelsPower storage: lithium-ion batteries


6.5 Risk assessmentWhen assessing risks of a complex system such as the Satellite VDES, severalaspects have to be considered. This has been done in accordance to the tem-plate for the Concept of Operations (ConOps) provided by NASA [11]. Therisks can be decomposed as:

• Development risks

• Operational risks

• Disposal risks

• Environmental risks

Have in mind that these are early identified risks, there may be many more thatthe author has not yet identified at this early stage of the project.

6.5.1 Identified risksDevelopment risks

- Risk of overestimating performance requirements of the system, which in-creases costs.- Risk of technical problems arising in the design process, such as when inte-grating subsystems and payload in the chosen satellite form factor.- Risk of scheduling delays, thus increasing costs.- Risk of postponed decision at WRC in November 2019 for allocation of fre-quencies for satellite VDES up- and downlink.- Risk of miscommunication between parties developing the space segmentsuch as the satellite provider, payload provider and service provider leading toerrors that are hard to change in the aftermath.

Operational risks

- Risk of failure at launch (around 90% of all launches are successful) [118].- Risk of failure of satellite electronics when in orbit due to SSE.- Risk of antennas, solar panels and other deploying mechanisms not deployingin space.- Risk of collisions within satellite constellation or with other satellites onsimilar inclination and altitude.- Risk of overestimating coverage, having to launch additional satellites.


Disposal risks

- Risk of satellites not re-entering within 25 years.- Risk of satellites not burning up in the atmosphere when re-entering, due tousing materials with a too high thermal conductivity.

Environmental risks

- Risk of generating orbital debris, which can lead to a cascade effect.- Risk of generating environmental footprint (fossil fuels) in production ofspecific subsystems in particular batteries and solar panels, which require pre-cious metals that are hard to extract without a significant footprint.- Risk of contamination of atmosphere at launch.

6.5.2 Analysed risksFollowing the approach by Larson [5] the risks that were the most critical couldbe classified as shown in Table 6.10 and later placed in a risk classificationmatrix, see Figure 6.6. A high likelihood is represented by a ≥ 10% risk ofoccurring, a high consequence means that the risk is devastating to an extentof endangering the whole system if it occurs.

If a risk is in the red spectrum it needs by furthest extent possible to bemitigated. For example risk number six could be mitigated by having severallaunches to establish the constellation, instead of a single one. If the risk is inthe orange spectrum it should be mitigated, but doesn’t necessarily have to be.Lastly, a risk in the green spectrum can be ignored.


Table 6.10: Showcases the six most critical development and operational risksfor the VDES system of those that have been identified.

Nr. Identified risk Likelihood Consequence

Low Moderate High Low Moderate High

1. Risk of postponeddecision at WRC 19. X X

2. Risk of satellites not re-enteringwithin 25 years. X X

3.

Risk of collision withinsatellite constellation orwith other satellites at

similar inclination and altitude.

X X

4.Risk of overestimating

coverage, having tolaunch additional satellites.

X X

5.Risk of

deployment mechanismsfailing in orbit.

X X

6. Risk of failure atlaunch. X X

Figure 6.6: Showcases a risk classification matrix over the identified risks.

Chapter 7

Discussion

To begin with, having studied the process of establishing a VDES satellitesystem I have come to the understanding of how such a process is very difficult.Where do you start and to what degree do one analyse each and every elementof the system?

After reading the article by Reid et al. [17] it was argued in section 2.4.3for why LEO was preferable over MEO and of most significance if one desiredto launch standardised small satellites such as Cubesats in order to minimizecosts. Taking this into consideration, it was decided by the author to proceedwith LEO and the CubeSat form factor to establish a LEO constellation forsatellite VDES.

Moreover, an analysis of the current market has been brought forward insection 5.1, where current AIS satellite systems on the market have been dis-sected, coming to the conclusion that at the time of writing 181 satellites in-orbit are using some type of the AIS system to receive signals from vesselscarrying AIS transponders. It has also been shown that more than 677 satel-lites with a AIS/VDES payload are expected to be placed in orbit within theupcoming years. In addition, one can distinguish a tendency that many timesit is a question of the same companies in circulation to provide satellite plat-forms within this domain, such as GomSpace, ÅAC Clyde Space and UTIASSFL. Especially GomSpace has been very active on the market and is not justproviding the satellite platform, but also often their own SDR platform for AISpurposes.

Furthermore, a more thorough analysis was done in Table 5.2 looking atsatellites using the CubeSat form factor to provide AIS/VDES services, whereit was noted that most have been launched very recently, within the last 4 years,all use solar power, all are located in LEO at an altitude that differ between 400-

84

CHAPTER 7. DISCUSSION 85

650 km and most commonly in sun synchronous orbit to optimize their powergeneration. The 3 U form factor was the most common, although, bigger formfactors have also been used. Means of communication was in general by S-band providing a few Mbps, even though X-band is an upcoming trend withmuch higher speeds as foreseen by Poghosyan and Golkar [15].

Another trend for CubeSats in constellations is the capability of ISL, whichas pointed out by Kocak and Browning [29] is a crucial need for real time datalatency and to minimize the number of satellites and ground stations that oth-erwise have to be in place to gain the same capability with a store forward, alsoknown as bent pipe satellite system. Instead, routing can be used to identifysatellites currently having ground stations within their FOV, in order to trans-fer the data to those satellites by ISL for downlink. Lastly, it should be addedthat most of the analysed CubeSats do not utilize any means of propulsion tomaintain their orbits, this is probably due to limited power supply.

The design concluded with a design budget allocated for a 6 U satelliteplatform, provided by ISIS, as can be seen in Section 6.4.1. The author wantsto highlight that it was a simplified budget due to limited information providedby the satellite providers on their websites. Especially ÅAC Clyde Space,which 6 U- and 3 U satellite platforms were of special interest (the later ifSaab’s SDR payload can be optimized power wise to around 15 W OAP), didnot provide any data in order to produce a design budget. ISIS 6 U platformwas the third best option and did provide enough data on its website for thedifferent subsystems to produce a budget. However, ISIS lacked both a powerstorage system and a power distribution and control system for its 6 U platform.Thus the author had to chose them from another provider, and as recommendedby ISIS on their website they were picked from GomSpace.

Moreover, pinpointing the launch segment, see section 4.5, and in partic-ular the launch vehicle, more data needs to be collected in order to make adefinite choice of how to cheapest launch the constellation in orbit, while stillfulfilling the orbital requirements. As for now the dedicated ride-share optionas well as launching as primary payload, by e.g the Electron launch vehicleprovided by Rocket Lab, should be further considered. On the other hand, Pig-gybacking (launching as secondary payload), although having proven to be arelative cheap option to get into orbit [119], tends to not give an exact enoughorbital geometry, especially considering that no propulsion system will be as-signed on the CubeSats.

Considering the ground segment, more work has to be made to optimizethe number of ground stations needed in order to decide upon a ground stationprovider or to build a designated ground station network for the system itself.

86 CHAPTER 7. DISCUSSION

Nevertheless, this is dependent on the routing of data from the satellite con-stellation, which yet has to be evaluated as well. Another question is if Saab’sbase stations, already employed around the world for terrestrial AIS/VDES,can function as relay mechanisms for the satellite segment and thus decreasethe number of ground stations needed for the system.

The constellation analysis that has been done, see Section 6.3, shows theminimum number of satellites needed to continuously cover the whole sur-face of Earth from an altitude of 400 km - 700 km in polar (or near-polar)orbits. This doesn’t account for any overlap between satellites’ FOVs and asthe Earth is not fully spherical perturbations will affect the orbit of each andevery satellite differently, which could create coverage gaps within the constel-lation. Initially, a maximum FOV of horizon to horizon (minimum elevationangle of ϵ = 0) was considered, this resulted in 32-50 satellites depending onaltitude. More realistic was instead to increase the minimum elevation angleto 10 degrees, as commonly done, which decreased the swath width and thusincreased the number of satellites needed to around three times the order forcontinuous global coverage.

To further explain the reasoning to increase the minimum elevation angle,it could be that it is restricted by the antenna chosen. This has to do withthe fact that low elevation angles require either a high gain antenna or a highpower transmitter to be used to not decrease the received power to an extentwhere it is indistinguishable from noise, see again equation 2.26. This is atrade-off and as we know power is scarce in small satellites. On the otherhand, a high gain directing antenna implies a smaller power beamwidth, theangle which indicates a reduction of half the direction gain (±3 dB), and asthe angle decreases the ”communication footprint” of the satellite decreases,such that directing the antenna or rotating the whole satellite might be neededto not limit the satellite’s coverage [120]. In the design of the satellite, theauthor chose to use a high gain antenna for the payload communication andas a result decrease the power needed for transmissions, however, as it alsomeans a smaller beamwidth, directing the antenna will be of more importanceto cover the whole FOV of the satellite.

A link- and a data budget have been conducted considering an orbital al-titude of 550 km. As for the link budget, one could notice that the uplinkperformed significantly better than the downlink for both cases, when keep-ing the gain constant and when having it change with the elevation angle. Thiswas true for both the carrier to noise density ratio, C

N0, and the energy per bit to

noise spectral density ratio, Eb

N0, as a higher value indicates better linking, see

Figures 6.3 and 6.4. Moreover, in Figure 6.3 the linking performs better with


increasing elevation angles, which is as expected since the free space path loss,Lfs, decreases, see equation 2.28, and the antenna gains are kept constant.

Considering Figure 6.4, only carrier to noise density ratio was evaluatedsince a MCS has not yet been decided and as they come with different datarates and bandwidths, it was the most practical approach to analyse the linkingagainst thresholds that could be expected for different ones, see ITU’s reportRep. ITU-R M.2435-0 [117] for more details on data rates and bandwidthsfor different MCS. Another observation is that the linking decreases for bothup- and downlink from an elevation angle of 40 degrees and onwards until 90degrees, this is mainly because of the antennas which have a maximum gain atminimum elevation angles. As for the satellite antenna, it is pointed towardsthe horizon at an offset nadir angle of around 66-67 degrees and that is whyit’s gain decreases until 90 degrees of elevation, when the nadir angle, (η = 0).Furthermore, a few thresholds for various modulation schemes retrieved from[117] are shown in Figure 6.4, where the downlink is below the threshold forthe modulation 8 Phase Shift Keying (8PSK) already at 80 degrees of eleva-tion, a similar trend is seen in the uplink although not before 85 degrees forthe modulation 16 Quadrature Amplitude Modulation (16QAM).

Having taken MCS into consideration, one could also desire to decrease theFOV in order to receive fewer VDES messages to each satellite. This would de-crease the risk of having data packet collisions occurring, something that mayhappen when too many vessels carrying AIS/VDES transponders are withina satellite’s FOV and transmitting simultaneously disturbing the decoding ofmessages in the satellite, as described by Kocak and Browning [29]. This iswhy Multiple Access (MA) will be used for the system [117], but whetherTime Division Multiple Access (TDMA) or Code Division Multiple Access(CDMA) will be used is unclear.

However, since it has not been the purpose of this work to analyse MCSand MA in more detail, it is left to future studies to decide what option couldbe preferable from a system point of view.

Regarding the data budget, Figure 6.5 showcases how the amount of datacollected during time of contact with the satellite for a single vessel changeswith the minimum elevation angle that is used. If a minimum elevation angleof ϵ = 10 degrees is used, around 105 Mb can be collected during one satellitepass when linking to only one vessel. This analysis was done using the freedata rate of 240 kb/s to simply the calculations and as a result of the MCS notyet being defined. This is in many ways an overestimation, obviously as therewill be many ships within the satellite’s FOV and some type MA will have to beused as a result, such that each vessel receives a specific place in the queue with

88 CHAPTER 7. DISCUSSION

a specific amount of time for data transfers, also known as timeslot, dependingon the importance of the data transfer. Furthermore, many modulation typeslisted in [117] have lower data rates.

How about Saab’s role in all this? In the development of the system thereare several scenarios, as identified by the author, in which Saab takes differentroles:

• Saab takes the role as payload provider leaving the development of thesystem to a service provider. It sells its payload, but has no control ofthe actual system.

• Saab becomes a service provider developing the whole constellation byitself, to have full control of the system and to sell the capabilities towhom it likes.

• Saab enters some kind of consortium where it acts as the payload provider,but cooperates with a satellite provider, a ground station provider and asatellite operator. It gains some control of the system development andoperations, hence has some control of who the users and customers are.

• Saab launches a few satellites with its own payload contributing to theecosystem of a constellation. Maybe some of its competitors join, suchas those listed earlier in subsection 3.2.1?

The recommendation of the author is that Saab proceeds to build and op-erate a satellite, carrying a SDR-platform for VDES purposes as payload, totest and evaluate its capabilities. Then by the additional gained informationSaab can take a more informed decision on regard of this matter.

Although, drawing from the result of this study, it seems as the market isalready crowded by service providers that offer satellite AIS services. Takingthis into consideration, maybe the most successful approach for Saab would beto take the role as payload provider to offer an SDR payload with the increasedcapability of VDES, in comparison to AIS, to sell to companies such as thoselisted in Table 5.1. By doing so, taking advantage of CubeSats short lifetime,making way for perhaps a never ending demand during the next 10-20 years.


7.1 Future workSome rather wide areas are open for future studies:

• Simulation of constellation, using for example software such as Soft-ware Tool Kit (STK), to provide a coverage analysis over time and studyhow perturbations affect it. When and where does coverage gaps occur?

• De-orbit simulation using STK or General Mission Analysis Tool (GMAT)in order to analyse whether all satellites re-enter within 25 years afterend-of-life.

• Investigate routing of the satellite segment. How is data efficiently sentthrough the ISL layer to forward the data as fast as possible to the user?

• Cost model analysis, who pays for the system? what payment streamscan be distinguished? Where is Saab located in the chain?

• Development plan, how does the project plan for the development of thesystem look? what activities are dependent? And when should differentactivities be done?

• Examine how data packet collisions can be treated considering that satel-lites have a much broader FOV than what the conventional AIS systemwas built for, what type of MA and MCS can be used to resolve thisissue?

• Future navigation possibilities using a satellite VDES system withoutlatency delay and with its own positioning system.

• Investigation of data enhancement to predict and showcase vessel routesin real time by some type of interface, e.g a web interface.

Chapter 8

Conclusion

To begin with, looking back at the research questions that this study advertisedto resolve:

• What requirements are put on the payload and the satellite respectivelyfor the satellite VDES system?

• Is it possible to use the existing SDR-platform developed at Saab aspayload for the purpose?

• How can global coverage be achieved by incorporating a satellite con-stellation called satellite VDES into the terrestrial VDES system?

It has been concluded that the payload’s physical parameters drive thesatellite platform. It should be possible to use a VDES transponder as a pay-load on a conventional 6 U CubeSat together with possibly one or more addi-tional payloads for similar Earth observing tasks.

Moreover, it is possible to use the existing payload developed at Saab forthe task, using a reduced size of 1.5 U, but without changes to power and mass.However if a 3 U CubeSat is desired to be used, the power consumption shouldbe limited to around 10-15 W orbit average power.

Lastly, global coverage can be achieved using a satellite constellation ofCubeSats in polar orbit at 550 km altitude with at least 91 satellites, consider-ing a minimum elevation angle of 10 degrees, to establish full global coverage,see Table6.4. The orbital altitude was chosen to generate as broad field of viewas possible, but to still de-orbit (naturally decay) within 25 years after end oflife strictly due to aerodynamic drag.

90

Appendix A

System requirements

91

92 APPENDIX A. SYSTEM REQUIREMENTS

Table A.1: System interface requirements in the satellite VDES system influ-enced by what has been done by Øystein et al. [10], as well as M. Bradburyet al. [8].


IR-01 System segmentsThe system shall be composed ofa space segment, a ground segmentand a launch segment.

IR-1.1 Space segmentThe space segment shall consistof a constellation of CubeSatscarrying payloads for VDES.

IR-1.2 Ground segment

The ground segmentshall consist of aGround station network,and a MCC.

IR-1.3 Launch segment

The launch segmentshall consist of alaunch system anda separation system.

IR-2 Satellite platformThe satellite platform shallhave a high frequency downlinkand uplink interface (space segment - ground segment).

IR-3 PayloadThe primary payload shallconsist of a SDR-platformproviding VDES.

IR-3.1 Payload subsystemsThe SDR platform shall supportAIS channels 1-4, ASM channels 1-2and VDE channels 1-12.

IR-3.1.1 Payload transmissionThe payload shall have a downlinkand uplink in thefrequency domain 156-163 MHz.

IR-4 Antenna systemThe satellite shallhave a radio antenna systemfor VHF and UHF transfers.

IR-4.1 Antenna The antenna shall be stowed atat launch and deployed in-orbit.

APPENDIX A. SYSTEM REQUIREMENTS 93

Table A.2: System functional requirements in the satellite VDES system in-fluenced by what has been done by Øystein et al. [10], as well as M. Bradburyet al. [8].


FR-1 ConstellationA constellation ofCubeSats shall be designedto provide continuous global coverage.

FR-1.1 Altitude Satellites in the constellation shallhave a common altitude between 400-700 km.

FR-1.2 Inclination Satellites in the constellationshall be of common inclination of 70-110 .

FR-3 ISL The space segment shall utilize ISL.

FR-4 GNSSThe satellites shall carry aGNSS receiver inorder to provide positional data.

FR-5 Satellite operationsThe ground segmentshall monitor and provide controlover the space segment.

FR-6 Redundancy

The system shall have redundantsatellites in-orbitin case of failureof one of the operational satellitesin the constellation.

FR-7 Modes of operation

The satellite shall have severalmodes of operation, including sun pointing,nadir pointing, inter satellite pointing, de-tumbling mode and non-operation mode.

FR-8 VDES The satellite shall providecontinuous VDES operations.

FR-9 Data handling The space segment shall performon-board processing.

FR-10 Encryption The data shall be encrypted toensure system integrity.

FR-11 User validationUsers shall be identified andverified when accessingdata from the system.


Table A.3: System performance requirements in the satellite VDES influencedby what has been done by Øystein et al. [10], as well as M. Bradbury et al. [8].


PR-1 Duty CycleThe payload shall handle a dutycycle of 90% foroperational service.

PR-2 Latency The system shall have a latencydelay of less than 2 s one-way.

PR-3 Data storage TBD.

PR-4 Down-link capacity

The satellite platform shallhave a feed downlinkcapacity of >10 Mbit/s betweenspace segment andground segment.

PR-5 Up-link capacity

The satellite platform shallhave a feed uplinkcapacity of >10 Mbit/sbetweens space segmentand ground segment.

PR-6 Attitude determinationThe satellite shall determineits attitude with an error of lessthan 1 degree.

PR-7 Attitude controlThe satellite shall controlits attitude with an error ofless than 3 degrees.

PR-9 Antenna gainThe antenna shallhave an antennagain between 5 - 10 dBi.

PR-10 Power generation The satellite platform shallgenerate more than 60 W.

PR-11 Energy storage The satellite shall be able tostore more than 80 Wh.

PR-12 Payload availability The payload shall allow for aavailability of 99.99%.

PR-13 Payload noise figure The payload shall have a noisefigure of <1.7 dB.

APPENDIX A. SYSTEM REQUIREMENTS 95

Table A.4: System constraint requirements in the satellite VDES system in-fluenced by what has been done by Øystein et al. [10], as well as M. Bradburyet al. [8].


CR-1 Lifetime

Each satellite in theconstellation shall haveat least 3 yearsoperational in-orbit lifetime.

CR-2 End-of-life Satellites shall re-enter within25 years after operational end of life.

CR-3 Satellite platform dimensionsThe satellite platform shallconsist of a 6 U standardisedCubeSat or lesser.

CR-3.1 Payload dimensionsThe VDES Payload shall notexceed the dimensionsspecified by the satellite platform.

CR-4 Satellite platform mass The satellite platform’s launchmass shall be less than 10 kg.

CR-4.1 Payload mass The payload mass includinghousing shall be less than 1.5 kg.

CR-4.2 Antenna mass The antenna mass shall be<0.5 kg.

CR-5.1 Payload power consumptionThe VDES payload shall have anaverage power consumptionof less than or equal to 20 W.

CR-5.2 Antenna power consumptionThe antenna shallhave a power consumptionof less than 4 W.

CR-6 Antenna size

The antenna systemshall be ableto accommodate onthe satellite platform.


Table A.5: System environmental requirements in the satellite VDES systeminfluenced by what has been done by Øystein et al. [10], as well as M. Brad-bury et al. [8].


ER-1 RadiationThe Space segment shall handlethe expected radiationfrom 3+ years in LEO without degradation.

ER-2 Thermal

The thermal system of thesatellite shall keepthe payload in therequired operational temperaturerange −20 - +60 Celsius.

ER-3 VibrationThe space segment shallhandle the mechanical loadsfrom an arbitrary launch system.

Appendix B

Link and data budgets

Table B.1: The link budget using constant antenna gains for downlink fromsatellite to subject (vessel).

ϵ [] Lfs [dBW] Pr [dBW] CN0

[dB Hz] EbN0

[dB]

0 145.28 −137.28 61.12 7.3210 141.82 −133.82 64.58 10.7820 138.87 −130.87 67.53 13.7230 136.58 −128.58 69.82 16.0240 134.83 −126.83 71.57 17.7750 133.53 −125.53 72.87 19.0760 132.58 −124.58 73.82 20.0270 131.94 −123.94 74.46 20.6680 131.57 −123.57 74.83 21.0390 131.45 −123.45 74.95 21.15

97

98 APPENDIX B. LINK AND DATA BUDGETS

Table B.2: The link budget using constant antenna gains for uplink from sub-ject (vessel) to satellite.

ϵ [] Lfs [dBW] Pr[dBW] CN0

[dBHz] EbN0

[dB]

0 145.01 −129.22 73.68 19.8710 141.54 −125.76 77.14 23.3320 138.60 −122.82 80.08 26.2830 136.30 −120.52 82.38 28.5840 134.56 −118.78 84.12 30.3250 133.25 −117.47 85.43 31.6260 132.31 −116.53 86.37 32.5770 131.67 −115, 89 87.01 33.2180 131.30 −115.51 87.39 33.5890 131.17 −115.39 87.51 33.71

Table B.3: The link budget for the satellite downlink to the subject (vessel),when the antenna gains of the satellite and ship are changing with the elevationangle as described in [117].


[dBHz]

0 145.28 −137.28 61.1210 141.82 −133.82 64.5820 138.87 −131.37 67.0330 136.56 −130.78 67,.6240 134.83 −130.93 67.4750 133.53 −132.53 65.8760 132.58 −134.98 63.4270 131.94 −138.24 60.1680 131.57 −146.77 51.6390 131.45 −159.95 38.45

APPENDIX B. LINK AND DATA BUDGETS 99

Table B.4: The link budget for the satellite uplink to the subject (vessel), whenthe antenna gains of the satellite and ship are changing with the elevation angleas described in [117].


[dBHz]

0 145.01 −129.22 73.6810 141.54 −125.76 77.1420 138.60 −123.32 79.5830 136.30 −122.72 80.1840 134.56 −122.88 80.0250 133.25 −124.47 78.4360 132.31 −126.93 75.9770 131.67 −130.19 72.7180 131.30 −138.71 64.1990 131.17 −151.89 51.01

Table B.5: The data budget shows how the time of satellite contact and theamount of data collected from a single vessel are dependent on the minimumelevation angle.

ϵ[] Tc [min] DA [Mb]0 11.24 161.9110 7.32 105.3620 4.94 71.2030 3.49 50.2240 2.52 36.2950 1.82 26.1960 1.27 18.2770 0.81 11.6080 0.39 5.6490 0 0

Appendix C

Concept of operations

100

Concept of Operations

This Concept of Operations (ConOps) annotated outline describes the type and sequence of information that should be contained in a ConOps, although the exact content and sequence will be a function of the type, size, and complexity of the project. The text in italics describes the type of information that would be provided in the associated subsection. Additional subsections should be added as necessary to fully describe the envisioned system.

1.0 Introduction

1.1 Project Description This section will provide a brief overview of the development activity and system context as delineated in the following two subsections.

1.1.1 Background Summarize the conditions that created the need for the new system. Provide the high-level mission goals and objective of the system operation. Provide the rationale for the development of the system.

See report: Background, AIS and the development of VDES.

1.1.2 Assumptions and Constraints State the basic assumptions and constraints in the development of the concept. For example, that some technology will be matured enough by the time the system is ready to be fielded, or that the system has to be provided by a certain date in order to accomplish the mission.

Assumptions:

- Spherical earth

- Circular orbits

- Keplerian orbits

- Non-rotating Earth

- Common inclination and altitude of constellation

- Average solar activity

- No third body perturbations

Constrains:

- Feasibility study

- Primary payload SDR-platform for VDES provided by Saab

- Follow standard provided by ITU and IALA

- CubeSat form factor

- LEO Orbit

- De-orbit within 25 years

1.2 Overview of the Envisioned System This section provides an executive summary overview of the envisioned system. A more detailed description will be provided in Section 3.0

1.2.1 Overview This subsection provides a high-level overview of the system and its operation. Pictorials, graphics, videos, models, or other means may be used to provide this basic understanding of the concept.

Architectural overview: See figures in report, Context diagram, System Architecture and WBS.

Operation overview: see figure Functional Flow Block Diagram in report.

1.2.2 System Scope This section gives an estimate of the size and complexity of the system. It defines the system’s external interfaces and enabling systems. It describes what the project will encompass and what will lie outside of the project’s development.

Internal interface: satellite segment – Satellites in constellation carrying VDES payload, ground

segment - ground stations and MCC and launch segment - launch vehicle and separation system.

External interface: Terrestrial VDES = Vessel segment + Shore segment in the VDES system.

Enabling systems: political decisions (World Radiocommunication Conference 2019 (WRC - 19) to

provide the frequencies), enforcement/mandatory VDES on all class A-vessels (IMO SOLAS).

2.0 Documents

2.1 Applicable Documents This section lists all the documents, models, standards or other material that are applicable and some or all of which will form part of the requirements of the project.

Standards: CalPoly CubeSat standard: form factor of satellite platform and separation system.

Space debris: the ESA approach, compendium of space debris mitigation standards adopted by

states and international organization.

2.2 Reference Documents This section provides supplemental information that might be useful in understanding the system or its scenarios.

- Space Norway: development and testing of satellite AIS and VDES.

- ExactEarth: development of satellite constellation to provide the service of global AIS.

- ESA mullighetsstudie : Satellite AIS (2009) .

- Wertz et al: SMAD 3rd edition (1999).

- Poghosyan et al: CubeSat evolution: Analyzing CubeSat capabilities for conducting science

missions (2017).

- Lazaro et al: VDES: an enabling technology for maritime communications (2018)

- Reid et al: Broadband LEO constellations for navigation (2018)

- Saab data sheet: Airborne R5A

3.0 Description of Envisioned System This section provides a more detailed description of the envisioned system and its operation as contained in the following subsections.

3.1 Needs, Goals and Objectives of Envisioned System This section describes the needs, goals, and objectives as expectations for the system capabilities, behavior, and operations. It may also point to a separate document or model that contains the current up-to-date agreed-to expectations.

Needs:

- Over the horizon capability ship-ship and ship-shore

- Instant latency (<2s) (real time)

- Global coverage

- Provide services outlined by ITU and IALA

Goal and objectives:

- To establish a satellite constellation that offers global continuous coverage of Earth using

VDES

- Enable AIS-uplink, ASM-uplink and VDE-up- and downlink

- Autonomous constellation

- Routing of VDES messages

3.2 Overview of System and Key Elements This section describes at a functional level the various elements that will make up the system, including the users and operators. These descriptions should be implementation free; that is, not specific to any implementation or design but rather a general description of what the system and its elements will be expected to do. Graphics, pictorials, videos, and models may be used to aid this description.

The system is referred to as the satellite VDES system and functions as on of three segments in the whole VDES system. The satellite VDES system is divided into three segments: Space segment, Launch segment and

Ground segment.

Space segment: Satellite platform, payload (SDR-platform) and satellite constellation.

Ground segment: Ground station network, MCC and perhaps data center.

Launch segment: Launch vehicle and separation/deployment system.

User segment: vessel segment and shore segment in the VDES system.

Users: Maritime authorities, shipping companies, mariners, others.

3.3 Interfaces This section describes the interfaces of the system with any other systems that are external to the project. It may also include high-level interfaces between the major envisioned elements of the system. Interfaces may include mechanical, electrical, human user/operator, fluid, radio frequency, data, or other types of interactions.

See Figure 3.2 and Figure 4.2, showcasing use-cases and interfaces in report. Example on interfaces are:

- IP (Ground segment - User segment (shore segment))

- Communications (Space segment – User segment (vessel segment)) by VDES on the maritime

band.

- Separation system (Space segment - Launch segment)

- Telemetry, tracking and command (Space segment - Ground segment)

- ISL (Space segment - Space segment)

3.4 Modes of Operations This section describes the various modes or configurations that the system may need in order to accomplish its intended purpose throughout its life cycle. This may include modes needed in the development of the system, such as for testing or training, as well as various modes that will be needed during it operational and disposal phases.

When system is in development:

-Testing: radiation testing, vacuum testing, EMC testing, vibrational load testing, and

temperature testing.

- Training (satellite operators, ground station operators).

When system is operational:

- VDES: transmitting and receiving

- Telemetry, tracking and command (up and down-link)

- Attitude determination and control

- On board processing

- Thermal control (passive)

- Power generation

- Power storage

- Power distribution

Pointing modes:

-Nadir pointing: the spacecraft is controlled w.r.t to the inertial reference frame for nadir pointing of ground target (vessels or ground stations). -Sun pointing: Spacecraft is controlled w.r.t the inertial reference frame for maximum solar input. -Inter Satellite pointing: the spacecraft is controlled w.r.t to the inertial reference frame to point to satellite in track or adjacent orbital plane. -De-tumbling mode: reduces spacecraft's angular rates at separation from launch vehicle as well as functions as a fallback safe mode during anomalies.

-Non-operation mode: satellite's attitude is uncontrolled, to for example save battery.

3.5 Proposed Capabilities This section describes the various capabilities that the envisioned system will provide. These capabilities cover the entire life cycle of the system’s operation, including special capabilities needed for the verification/validation of the system, its capabilities during its intended operations, and any special capabilities needed during the decommissioning or disposal process.

-The space segment withstands the vibrational loads at launch.

- The space segment tolerates the space environment in LEO.

-The space segment has the capability to de-orbit within 25 years.

- Capability of Inter satellite communication in space segment.

- Capability for ground segment to validate that all subsystems of each individual satellite is

working correctly.

- Capability for each respective satellite in the space segment to receive and process thousands

of VDES messages instantaneously while transmitting VDES messages.

- Capability for on board processing of data within space segment.

- Capability for space segment to unpack and identify key VDES messages in-orbit.

- Capability for space segment to route VDES messages efficiently within satellite network to

deliver it to the right user.

- Capability for ground segment to upload software updates to space segment in-orbit

enhanced by the SDR-platform.

4.0 Physical Environment This section should describe the environment that the system will be expected to perform in throughout its life cycle, including integration, tests, and transportation. This may include expected and off-nominal temperatures, pressures, radiation, winds, and other atmospheric, space, or aquatic conditions. A description of whether the system needs to operate, tolerate with degraded performance, or just survive in these conditions should be noted.

Integration, testing and transportation

Integration of space segment: Payload and subsystems to satellite structure.

Integration between space segment and launch segment: Satellite to be integrated to

separation system and fixed in launch vehicle.

Testing of various subsystems for space environmental effects, as earlier mentioned.

Transportation: from satellite development facilities to launch facilities. Expected space environmental effects:

Space segment needs to tolerate without degraded performance vibrational loads at launch.

Space segment needs to tolerate with degraded performance: radiation of 3krad for 5+ years.

Space segment needs to operate with thermal nominal temperatures of -20 - +60 degrees

Celsius.

Space segment needs to operate in a vacuum environment.

Space segment needs to operate with changing solar activity.

5.0 Support Environment This section describes how the envisioned system will be supported after being fielded. This includes how operational planning will be performed and how commanding or other uploads will be determined and provided, as required. Discussions may include how the envisioned system would be maintained, repaired, replaced, it’s sparing philosophy, and how future upgrades may be performed. It may also include assumptions on the level of continued support from the design teams.

Maintenance:

- Ground segment will have active maintenance.

- Space segment cannot be given physical maintenance of the hardware, but instead

maintenance can be done for the software, through software uploads and orbit

corrections through attitude control.

Reparations:

- Ground segment: physical and non-physical reparations can be done.

- Space segment - non-physical reparations can be done.

- Launch segment: reparations can only be done pre-operations.

Replacement:

- Space segment: satellites will be replaced every five years, and active spares will be in

orbit if satellites fail, to quickly re-establish global coverage.

- Ground segment: No expected replacement during the horizon of the mission.

- Launch segment: replacement of launch vehicle for each individual launch.

Upgrades:

- Space segment: Software updates regularly to satellite subsystems and especially to

payload taking advantage of the potential of the SDR-platform. Hardware updates may

play part to next generation of satellites replacing the first after 5+ years.

- Launch segment: No updates done.

- Ground segment: software and physical updates may take place.

6.0 Operational Scenarios, Use Cases and/or Design Reference Missions This section takes key scenarios, use cases, or DRM and discusses what the envisioned system provides or how it functions throughout that single-thread timeline. The number of scenarios, use cases, or DRMs discussed should cover both nominal and off-nominal conditions and cover all expected functions and capabilities. A good practice is to label each of these scenarios to facilitate requirements traceability; e.g., [DRM-0100], [DRM- 0200], etc.

6.1 Nominal Conditions These scenarios, use cases, or DRMs cover how the envisioned system will operate under normal circumstances where there are no problems or anomalies taking place.

Nominal conditions

- All satellites operative, full global continuous coverage.

- All ground stations operational.

- All data communication functional.

- All operational modes functional.

Reference documents

ITU-M 2090: requirements regarding VDES interface.

ESA mullighetstudie for et europeisk satellittbasert AIS system: system requirements for the

whole system.

CalPoly- CubeSat Design Specification: form factor for satellite size and mass.

Use cases:

- Positioning

- SAR:

- Transmission of FAL

- Broadcast of MSI

- Route exchange

- Download of updated digital publications

6.2 Off-Nominal Conditions These scenarios cover cases where some condition has occurred that will need the system to perform in a way that is different from normal. This would cover failures, low performance, unexpected environmental conditions, or operator errors. These scenarios should reveal any additional capabilities or safeguards that are needed in the system.

- Several satellites malfunctioning: not achieving global coverage - must set operation

effectively to reduce revisit time.

- Data packet collisions in satellite: A high efficiency random access protocol for satellites.

- Difficult weather conditions: empirical path losses – increased minimum elevation angle

or antenna gain.

- Satellite starts tumbling: de-tumbling mode.

- Solar eclipse: Prioritize payload operations.

7.0 Impact Considerations This section describes the potential impacts, both positive and negative, on the environment and other areas.

7.1 Environmental Impacts Describes how the envisioned system could impact the environment of the local area, state, country, worldwide, space, and other planetary bodies as appropriate for the systems intended purpose. This includes the

possibility of the generation of any orbital debris, potential contamination of other planetary bodies or atmosphere, and generation of hazardous wastes that will need disposal on earth and other factors. Impacts should cover the entire life cycle of the system from development through disposal. .

- Possibility of generating orbital debris, which can lead to a cascade effect.

- Possibility of re-entering satellites not burning up in the atmosphere.

- Environmental footprint (fossil fuels) in production of specific subsystems in particular

the batteries and the solar panels, which require precious metals that are hard to

extract without a significant footprint.

- Contamination of atmosphere at launch.

7.2 Organizational Impacts Describes how the envisioned system could impact existing or future organizational aspects. This would include the need for hiring specialists or operators, specialized or widespread training or retraining, and use of multiple organizations.

- Hiring of satellite provider to build the satellite platform and to integrate the

subsystems and payload onto the platform.

- Hiring of satellite operators to operate the satellites.

- Multiple organizations involved in the standardization of VDES: ITU, IALA, ESA, EMSA

etc.

- Decision not given at WRC - 19 to allocate frequencies for downlink from satellite on

maritime band (ITU).

- Decision on making it mandatory to carry VDES transponder on all class-A vessels (IMO,

SOLAS).

7.3 Scientific/Technical Impacts This subsection describes the anticipated scientific or technical impact of a successful mission or deployment, what scientific questions will be answered, what knowledge gaps will be filled, and what services will be provided. If the purpose of this system is to improve operations or logistics instead of science, describe the anticipated impact of the system in those terms.

Scientific improvements on:

- How to avoid data packet collisions in satellites with a much broader field of view than

what the conventional AIS system was built for.

- How to enhance the routing of VDES messages in a space segment using inter satellite

link.

- Possibility for autonomous navigation of vessels.

- Making it possible to provide route exchange between vessel- and shore segment in the

VDES system globally.

8.0 Risks and Potential Issues This section describes any risks and potential issues associated with the development, operations or disposal of the envisioned system. Also includes concerns/risks with the project schedule, staffing support, or implementation approach. Allocate subsections as needed for each risk or issue consideration. Pay special attention to closeout issues at the end of the project.

Development:

- Increasing costs by overestimating performance requirements of the system.

- Technical problems arise in design process, such as fitting all subsystems in the chosen

satellite form factor.

- Scheduling delays happen, which risk to increase costs.

- WRC 19 - allocation of frequencies for up- and downlink is postponed.

- Miscommunication between parties developing the space segment such as the satellite

provider, payload provider and launch provider leading to errors that are hard to change

in aftermath.

Operations:

- There is a risk of failure at launch (around 90% of all launches are successful).

- Risk of failure of satellites when in orbit due to radiation degradation.

- Risk of antennas, solar arrays and other deployment mechanisms not deploying in

space.

- Risk of collisions within satellite constellation or with other satellites on similar

inclination and altitude.

- Risk of overestimating coverage, having to launch additional satellites.

Disposal:

- Risk of no re-entry after 25 years.

- Risk of satellite not burning up when re-entering, due to using materials with a too high

thermal conductivity.

Appendix D

MATLAB code

1 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

2 % C a l c u l a t i o n program to e s t i m a t e t h e minimumnumber o f s a t e l l i t e s needed f o r a LEOc o n s t e l l a t i o n

3

4 % Date : 2019−06−105 % Author : J u l i a n G r u j i c i c6

7 % Approx ima t i ons :8 % A s p h e r i c a l non−r o t a t i n g Ea r t h .9 % C i r c u l a r s a t e l l i t e o r b i t s .

10 % A c o n s t e l l a t i o n o f P o l a r o r b i t s i s c o n s i d e r e d .11

12 %% Hor izon t o h o r i z o n maximum swath wid the l e v a t i o n minimum : e p s i l o n =0

13

14 c l e a r a l l15 c l c16 c l o s e a l l17

18 % C o n s t a n t s19 R=6371 ; % Mean Rad ius o f Ea r t h [km] Source : ”

Fundamen ta l s o f As t rodynamics ” [K. F Wakker ]20 mu=398600 .4418 ; % G r a v i t a t i o n a l p a r am e t e r [km^3 / s

109

110 APPENDIX D. MATLAB CODE

^ 2 ] , Source : ” Fundamen ta l s o f As t rodynamics ” [K.F Wakker ]

21 Re=6378 ; % Radius o f Ea r t h where i t i s a t i t sl a r g e s t a t t h e e q u a t o r [km] , Source : [ Nasa ] ,u r l : h t t p s : / / imag ine . g s f c . na sa . gov / f e a t u r e s /cosmic / e a r t h _ i n f o . h tml

22

23 % O r b i t a l speed and p e r i o d f o r d i f f e r e n t o r b i t a la l t i t u d e s .

24 H= 4 0 0 : 2 5 : 7 0 0 ; % O r b i t a l a l t i t u d e o f s a t e l l i t e [km]25 V= s q r t (mu . / ( H+R) ) ; % V e l o c i t y [km / s ]26 T= (2 . ∗ p i . ∗ s q r t ( (H+R) . ^ 3 . / mu) ) / 6 0 ; % O r b i t a l p e r i o d

[ min ]27

28 % Ea r t h cove r age a n g l e s and d i s t a n c e s .29 lambda_0= acos ( Re . / ( Re+H) ) ; % Maximum Ea r t h c e n t r a l

a ng l e [ r ad ]30 rho= a s i n ( Re . / ( Re+H) ) ; % Angula r r a d i u s o f e a r t h [

r ad ]31 swath =2∗ lambda_0 ; % The swath a t t h e t r u e h o r i z o n [

r ad ]32 e t a = p i /2− lambda_0 ; % The n a d i r a ng l e a t t h e t r u e

h o r i z o n [ r ad ]33 swa th_wid th =2 .∗ lambda_0 . ∗ Re ; % Swath_width [km]34 D_0=Re . / t a n ( rho ) ; % The s l a n t r ange a t t h e h o r i z o n

which i s t h e d i s t a n c e from t h e s a t e l l i t e t o t h et r u e h o r i z o n [km]

35

36 % Number o f s a t e l l i t e s i n c o n s t e l l a t i o n37 o r b i t a l _ p l a n e s = c e i l ( ( 2 ∗ p i ) . / ( 2 ∗ swath ) ) ; % The

number o f p l a n e s needed around t h e S p h e r i c a lEa r t h [−]

38 s a t e l l i t e s _ i n _ p l a n e = c e i l ( ( ( 2 ∗ p i ) . / ( swath ) ) ) ; % Thenumber o f s a t e l l i t e s needed i n each o r b i t a l

p l a n e [−]39 s a t e l l i t e s = o r b i t a l _ p l a n e s . ∗ s a t e l l i t e s _ i n _ p l a n e ; %

T o t a l number o f s a t e l l i t e s i n c o n s t e l l a t i o n [−]40

41 t a b l e (H’ , s a t e l l i t e s ’ ,V’ , T ’ , D_0 ’ ) % P r i n t i n g t a b l e

APPENDIX D. MATLAB CODE 111

showcas ing o r b i t a l a l t i t u d e , number o f s a t e l l i t e, o r b i t a l speed , o r b i t a l p e r i o d and s l a n t r angewhen c o n s i d e r i n g t h e t r u e h o r i z o n

42

43 %% Changing t h e e l e v a t i o n minimum ang l e betweene p s i l o n = 1 :10

44


49 % C o n s t a n t s50 R=6371 ; % Mean Rad ius o f Ea r t h [km] Source : ”

Fundamen ta l s o f As t rodynamics ” [K. F Wakker ]51 mu=398600 .4418 ; % G r a v i t a t i o n a l p a r am e t e r [km^3 / s

^ 2 ] , Source : ” Fundamen ta l s o f As t rodynamics ” [K.F Wakker ]

52 Re=6378 ; % Radius o f Ea r t h where i t i s a t i t sl a r g e s t a t t h e e q u a t o r [km] , Source : [ Nasa ] ,u r l : h t t p s : / / imag ine . g s f c . na sa . gov / f e a t u r e s /cosmic / e a r t h _ i n f o . h tml

53

54 % O r b i t a l speed and p e r i o d f o r d i f f e r e n t o r b i t a la l t i t u d e s .

55 H= 4 0 0 : 1 0 : 8 0 0 ; % O r b i t a l a l t i t u d e o f s a t e l l i t e [km]56 V= s q r t (mu . / ( H+R) ) ; % O r b i t a l V e l o c i t y [km / s ]57 T= ( 2 . ∗ p i . ∗ s q r t ( (H+R) . ^ 3 . / mu) ) / 6 0 ; % O r b i t a l p e r i o d

[ min ]58 rho= a s i n ( Re . / ( R+H) ) ; % Angula r r a d i u s o f e a r t h [ r ad

]59 f o r i =0:1060 e l e v a t i o n p r i n t i +1 = s p r i n t f ( ’ e l e v a t i o n ang l e

%d \ \ c i r c ’ , i ) ; % For−l oop ove r d i f f e r e n ta n g l e s o f e l e v a t i o n minimum

61 e p s i l o n = deg2rad ( i ) ; % Trans fo rm e l e v a t i o n ang l efrom d e g r e e s i n t o r a d i a n s [ r ad ]

62 e t a = a s i n ( cos ( e p s i l o n ) . ∗ s i n ( rho ) ) ; % The n a d i ra ng l e a t t h e s p e c i f i c e l e v a t i o n ang l e [ r ad ]

63 lambda= p i /2− e t a−e p s i l o n ; % The Ea r t h c e n t r e


ang l e f o r t h e s p e c i f i c e l e v a t i o n minimumang l e [ r ad ]

64 swath =2∗ lambda ; % The swath ang l e [ r ad ]65 swa th_wid th =2 .∗ lambda . ∗ Re ; % Swath wid th [km]66 D=Re . ∗ ( s i n ( lambda ) . / s i n ( e t a ) ) ; % S l a n t r ange

which i s t h e d i s t a n c e from t h e s a t e l l i t e t ot h e ground t a r g e t l o c a t e d a t t h e e l e v a t i o nminimum ang l e [km]

67 o r b i t a l _ p l a n e s = c e i l ( ( 2 ∗ p i ) . / ( 2 ∗ swath ) ) ; % Thenumber o f p l a n e s needed around t h e S p h e r i c a l

Ea r t h [−]68 s a t e l l i t e s _ i n _ p l a n e = c e i l ( ( ( 2 ∗ p i ) . / ( swath ) ) ) ; %

The number o f s a t e l l i t e s needed i n eacho r b i t a l p l a n e [−]

69 s a t e l l i t e s = o r b i t a l _ p l a n e s . ∗ s a t e l l i t e s _ i n _ p l a n e ;% T o t a l number o f s a t e l l i t e s i n

c o n s t e l l a t i o n [−]70

71 f i g u r e ( 1 )72 p l o t (H, swath_wid th , ’− ’ , ’ c o l o r ’ , r and ( 1 , 3 ) , ’

L inew id th ’ , 2 ) % P l o t t i n g t h e o r b i t a la l t i t u d e vs swath wid th

73 y l a b e l ( ’ Swath wid th [km] ’ )74 x l a b e l ( ’ A l t i t u d e [km] ’ )75 ho ld on76 end77 l e g end ( e l e v a t i o n p r i n t )78 t i t l e ( ’ A l t i t u d e vs . swath width , when a l t e r i n g t h e

minimum e l e v a t i o n ang l e ’ )79 g r i d on80

81 t a b l e (H’ , s a t e l l i t e s ’ , o r b i t a l _ p l a n e s ’ ,s a t e l l i t e s _ i n _ p l a n e ’ ,D’ ) % P r i n t i n g t a b l eshowcas ing o r b i t a l a l t i t u d e , number o f s a t e l l i t e, o r b i t a l speed , o r b i t a l p e r i o d and s l a n t r angewhen c o n s i d e r i n g e l e v a t i o n minimum of 10 d e g r e e s

1 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%


2 % C a l c u l a t i o n o f r e q u i r e d i n c l i n a t i o n t o have a SunSynchronous O r b i t (SSO)

3

4 % Author : J u l i a n G r u j i c i c5 % Date 2019−07−036

7 % Approx ima t i ons :8 % S p h e r i c a l E a r t h9 % C i r c u l a r o r b i t s

10

11 c l c12 c l e a r a l l13 c l o s e a l l14

15 H= ( 4 0 0 : 1 0 : 7 0 0 ) ; %A l t i t u d e [km]16 RE=6371; % Mean r a d i u s o f Ea r t h [km]17 i = rad2deg ( acos ( −0 . 098916 .∗ ( (H+RE) . / RE) . ^ ( 7 / 2 ) ) ) ;%

I n c l i n a t i o n [ deg ]18 p l o t (H, i ) %P l o t t i n g f i g u r e19 x l a b e l ( ’ A l t i t u d e [km] ’ )20 y l a b e l ( ’ I n c l i n a t i o n [ \ c i r c ] ’ )21 t i t l e ( ’ I n c l i n a t i o n f o r C i r c u l a r SSO vs A l t i t u d e ’ )22 t a b l e (H’ , i ’ )

1 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

2

3 % C a l c u l a t i o n program to e s t i m a t e Link budge t4


8 % I d e a l c a s e : No a t m o s p h e r i c l o s s e s9 % No c a b l e l o s s e s

10 % No t r a n s m i s s i o n o r r e c e i v e r l o s s e s11

12 %% Downlink s a t−s h i p %%%%


13

14 c l c ;15 c l e a r a l l16 c l o s e a l l17

18 %C o n s t a n t s19 Re=6371 e3 ; % Mean r a d i u s o f Ea r t h [m]20 H=550 e3 ; % O r b i t a l t i t u d e [m]21 k =1.380649∗10^( −23) ; %Boltzmann ’ s Con s t a n t [m^2 kg

s ^(−2) K^(−1) ]22 k_dB=10∗ log10 ( k ) ; %Boltzmann ’ s c o n s t a n t [ dB ]23 Gt =8;% Gain o f Yagi−Uda an t enna [ dBi ] , a s g i ven i n

d e s i g n budge t24 Gr =3; % Gain o f s h i p an t e nna [ dBi ] , a s g i ven i n t e c h

r ep . ITU−R M.2435−025 Pt =1; % Power o f s a t e l l i t e t r a n s m i t t e r [W] , from

d e s i g n budge t26 f0 =162∗10^6; % C a r r i e r f r e qu en cy f o r Downlink [ Hz ] ,

a s g i ven i n t e c h r ep . ITU−R M.2435−027 c =299792458; % speed of l i g h t [m/ s ]28 Lp =3; % C i r c u l a r p o l a r i z a t i o n l o s s e s [ dB ] , a s

g i ven i n t e c h r ep . ITU−R M.2435−029 Ts = 3 0 . 2 ; % System n o i s e t e m p e r a t u r e a t s h i p

r e c e i v e r [ dBk ] , a s g i ven i n IALA’ s r e p o r tG u i d e l i n e G1139 , Annex D

30 B_R=240∗10^3; % B i t r a t e [ b / s ] , r e t i r e v e d from ITU−R’ s r e p o r t M.2092−0

31

32 %Minimum e l e v a t i o n ang l e and s l a n t r a n g e s33 e l d eg = l i n s p a c e ( 0 , 90 , 1000 ) ; %make a v e c t o r o f

e l e v a t i o n a n g l e s34 e l = deg2rad ( e l d eg ) ; % from d e g r e e s t o r a d i a n s [ r ad ]35 D=Re . ∗ cos ( p i /2+ e l ) + s q r t ( ( ( Re+H) . ^ 2 ) −(Re . ^ 2 ) +(Re . ^2∗

cos ( e l + p i / 2 ) . ^ 2 ) ) ; % s l a n t r ange [m]36

37 Pt_dbw =10∗ log10 ( P t ) ; % from Watt t o Dec ibe l−Watt [dBw]

38 EIRP=Pt_dbw+Gt ; % E n f f e c t i v e i s o t r p o i c r a d i a t e dPower [dBw]


39

40 l =c / f0 ; %wave l eng th [m]41 Lfs =20.∗ log10 ( ( 4 . ∗ p i . ∗D) . / l ) ; % Free space pa t h

l o s s e s [ dB ]42

43 Pr=EIRP+Gr−Lfs−Lp ; % Rece i v e r power [dBw]44

45 C_N0=Pr−Ts−k_dB ;% C a r r i e r t o n o i s e d e n s i t y r a t i o [dBHz ]

46

47 R_bit_dB =10∗ log10 (B_R) ; % I n t o Dec i b e l [ dBHz ]48 Eb_N0=C_N0−R_bit_dB ; % Energy pe r b i t t o n o i s e

power s p e c t r a l d e n s i t y r a t i o [ dB ]49

50 %P l o t t i n g51 f i g u r e ( 1 )52 s u b p l o t ( 2 , 1 , 1 )53 p l o t ( r ad2deg ( e l ) ,C_N0 , ’−−r ’ , ’ LineWidth ’ , 2 )54 y l a b e l ( ’C / N_0 [ dBHz ] ’ )55 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )56 g r i d on57 t i t l e ( ’ C a r r i e r t o d e n s i t y r a t i o vs e l v a t i o n ang l e ’ )58 ho ld on59

60 s u b p l o t ( 2 , 1 , 2 )61 p l o t ( r ad2deg ( e l ) , Eb_N0 , ’−−r ’ , ’ LineWidth ’ , 2 )62 y l a b e l ( ’Eb / N_0 [ dB ] ’ )63 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )64 g r i d on65 t i t l e ( ’ Energy pe r b i t t o n o i s e s p e c t r a l d e n s i t y

r a t i o vs . e l v a t i o n ang l e ’ )66 l e g end ( ’ Downlink : E_b / N_0 ’ )67 ho ld on68 %% Upl ink budge t sh ip−s a t69

70 c l c71 c l e a r a l l72

73 %C o n s t a n t s


74 Re=6371 e3 ; % Mean r a d i u s o f Ea r t h [m]75 H=550 e3 ; % O r b i t a l t i t u d e [m]76 k =1.380649∗10^( −23) ; %Boltzmann ’ s Con s t a n t [m^2 kg

s ^(−2) K^(−1) ]77 k_dB=10∗ log10 ( k ) ; %Boltzmann ’ s c o n s t a n t [ dB ]78 Gt =3;% Gain o f s h i p an t e nna [ dBi ] , a s g i ven i n t e c h

r ep . ITU−R M.2435−079 Gr =8; % Gain o f Yagi−Uda an t enna [ dBi ] , from d e s i g n

budge t80 Pt =6; %Power o f s h i p t r a n s m i t t e r [W] , a s g i ven i n

t e c h r ep . ITU−R M.2435−081 f0 =157∗10^6; % C a r r i e r f r e qu en cy f o r Upl ink [ Hz ] , 082 c =299792458; % speed of l i g h t [m/ s ]83 Lp =3; % C i r c u l a r p o l a r i z a t i o n l o s s e s [ dB ] , a s

g i ven i n t e c h r ep . ITU−R M.2435−084 Ts = 2 5 . 7 ; % System n o i s e t e m p e r a t u r e [ dBk ] a t

s a t e l l i t e r e c e i v e r , a s g i ven i n IALA’ s r e p o r tG u i d e l i n e G1139 , Annex D

85 B_R=240∗10^3; % B i t r a t e [ b / s ] , r e t i r e v e d from ITU−R’ s r e p o r t M.2092−0

86






94





100 C_N0=Pr−Ts−k_dB ;% C a r r i e r t o n o i s e d e n s i t y r a t i o [dBHz ]

101

102 R_bit_dB =10∗ log10 (B_R) ; % I n t o Dec i b e l [ dBHz ]103 Eb_N0=C_N0−R_bit_dB ; % Energy pe r b i t t o n o i s e

power s p e c t r a l d e n s i t y r a t i o [ dB ]104

105

106 %P l o t t i n g107 s u b p l o t ( 2 , 1 , 1 )108 p l o t ( r ad2deg ( e l ) ,C_N0 , ’−−g ’ , ’ LineWidth ’ , 2 )109 y l a b e l ( ’C / N_0 [ dBHz ] ’ )110 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )111 g r i d on112 t i t l e ( ’ C a r r i e r t o n o i s e d e n s i t y r a t i o vs . e l v a t i o n

ang l e ’ )113 l e g end ( ’ Downlink : C / N_0 ’ , ’ Upl ink : C / N_0 ’ , ’ Loc a t i o n ’

, ’ Sou t hEa s t ’ )114

115 s u b p l o t ( 2 , 1 , 2 )116 p l o t ( r ad2deg ( e l ) , Eb_N0 , ’−−g ’ , ’ LineWidth ’ , 2 )117 y l a b e l ( ’Eb / N_0 [ dB ] ’ )118 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )119 g r i d on120 t i t l e ( ’ Energy pe r b i t t o n o i s e s p e c t r a l d e n s i t y

r a t i o vs . e l v a t i o n ang l e ’ )121 l e g end ( ’ Downlink : E_b / N_0 ’ , ’ Upl ink : E_b / N_0 ’ , ’

Loc a t i o n ’ , ’ Sou t hEa s t ’ )

1 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

2

3 % C a l c u l a t i o n program to e s t i m a t e Link budge t w i thchang ing ga i n

4



8 % I d e a l c a s e : No a t m o s p h e r i c l o s s e s9 % No c a b l e l o s s e s

10 % No t r a n s m i s s i o n o r r e c e i v e r l o s s e s11

12 %% Upl ink budge t sh ip−s a t13 c l c14 c l e a r a l l15 c l o s e a l l16


s ^(−2) K^(−1) ]21 k_dB=10∗ log10 ( k ) ; %Boltzmann ’ s c o n s t a n t [ dB ]22 Gr =[8 8 8 7 . 8 6 . 9 5 . 5 3 . 6 0 . 7 −2.2 −5.5] ;% Gain o f

Yagi−Uda an t enna [ dBi ] , from d e s i g n budge t23 Gt =[3 3 2 . 5 1 0 −1.5 −3 −4 −10 −20]; % Gain o f s h i p

an t e nna [ dBi ] , a s g i ven i n t e c h r ep . ITU−R M.2435−0

24 Pt =6; %Power o f s h i p t r a n s m i t t e r [W] , a s g i ven i nt e c h r ep . ITU−R M.2435−0

25 f0 =157∗10^6; % C a r r i e r f r e qu en cy f o r Upl ink [ Hz ] , 026 c =299792458; % speed of l i g h t [m/ s ]27 Lp =3; % C i r c u l a r p o l a r i z a t i o n l o s s e s [ dB ] , a s

g i ven i n t e c h r ep . ITU−R M.2435−028 Ts = 2 5 . 7 ; % System n o i s e t e m p e r a t u r e [ dBk ] a t

s a t e l l i t e r e c e i v e r , a s g i ven i n IALA’ s r e p o r tG u i d e l i n e G1139 , Annex D

29

30 %Minimum e l e v a t i o n ang l e and s l a n t r a n g e s31 maxel =90; % maximum e l e v a t i o n ang l e [ deg ]32 e l d eg =0 : 10 : maxel ; %make a v e c t o r o f e l e v a t i o n

a n g l e s33 e l = deg2rad ( e l d eg ) ; % from d e g r e e s t o r a d i a n s [ r ad ]34 D=Re . ∗ cos ( p i /2+ e l ) + s q r t ( ( ( Re+H) . ^ 2 ) −(Re . ^ 2 ) +(Re . ^2∗


36 Pt_dbw =10∗ log10 ( P t ) ; % from Watt t o Dec ibe l−Watt [


dBw]37 EIRP=Pt_dbw+Gt ; % E n f f e c t i v e i s o t r p o i c r a d i a t e d

Power [dBw]38




44 C_N0=Pr−Ts−k_dB ;% C a r r i e r t o n o i s e d e n s i t y r a t i o [dB−Hz ]

45

46 %T h r e s h o l d s f o r d i f f e r e n t modu l a t i o n s as g iven i nt e c h r ep . ITU−R M.2435−0

47 QPSK_CDMA_C_N0= 3 1 . 7 ;% [ dBHz ]48 pi4_QPSK_C_N0 = 4 9 . 2 ; %[ dBHz ]49 PSK_C_N0 = 5 3 . 3 ; %[ dBHz ]50 QAM_C_N0= 5 7 . 5 ; %[ dBHz ]51

52 %P l o t t i n g53 s u b p l o t ( 2 , 1 , 1 )54 p l o t ( r ad2deg ( e l ) ,C_N0 , ’−−g ’ , ’ LineWidth ’ , 2 )55 y l a b e l ( ’C / N_0 [ dBHz ] ’ )56 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )57 g r i d on58 t i t l e ( ’ Upl ink : c a r r i e r t o n o i s e d e n s i t y r a t i o vs .

e l v a t i o n ang l e ’ )59 ho ld on60 p l o t ( r ad2deg ( e l ) ,QPSK_CDMA_C_N0. ∗ ones ( s i z e ( e l ) ) , ’−c

’ , ’ LineWidth ’ , 2 )61 p l o t ( r ad2deg ( e l ) , pi4_QPSK_C_N0 . ∗ ones ( s i z e ( e l ) ) , ’ b ’ ,

’ LineWidth ’ , 2 )62 p l o t ( r ad2deg ( e l ) ,PSK_C_N0 . ∗ ones ( s i z e ( e l ) ) , ’m’ , ’

LineWidth ’ , 2 )63 p l o t ( r ad2deg ( e l ) ,QAM_C_N0. ∗ ones ( s i z e ( e l ) ) , ’ k ’ , ’

LineWidth ’ , 2 )64 l e g end ( ’C / N_0 ’ , ’QPSK/CDMA’ , ’ \ p i / 4 QPSK ’ , ’ 8PSK ’ , ’ 16

QAM’ , ’ Loc a t i o n ’ , ’ No r t hEa s t ’ )


65

66

67 %% Downlink budge t s a t−s h i p %68


s ^(−2) K^(−1) ]73 k_dB=10∗ log10 ( k ) ; %Boltzmann ’ s c o n s t a n t [ dB ]74 Gt =[8 8 8 7 . 8 6 . 9 5 . 5 3 . 6 0 . 7 −2.2 −5.5] ;% Gain o f

Yagi−Uda an t enna [ dBi ] , from d e s i g n budge t75 Gr =[3 3 2 . 5 1 0 −1.5 −3 −4 −10 −20]; % Gain o f s h i p

an t e nna [ dBi ] , a s g i ven i n t e c h r ep . ITU−R M.2435−0

76 Pt =1; % Power o f s a t e l l i t e t r a n s m i t t e r [W] , fromd e s i g n budge t

77 f0 =162∗10^6; % C a r r i e r f r e qu en cy f o r Downlink [ Hz ] ,a s g i ven i n t e c h r ep . ITU−R M.2435−0

78 c =299792458; % speed of l i g h t [m/ s ]79 Lp =3; % C i r c u l a r p o l a r i z a t i o n l o s s e s [ dB ] , a s

g i ven i n t e c h r ep . ITU−R M.2435−080 Ts = 3 0 . 2 ; % System n o i s e t e m p e r a t u r e a t s h i p

r e c e i v e r [ dBk ] , a s g i ven i n IALA’ s r e p o r tG u i d e l i n e G1139 , Annex D

81

82 %Minimum e l e v a t i o n ang l e and s l a n t r a n g e s83 maxel =90; % maximum e l e v a t i o n ang l e [ deg ]84 e l d eg =0 : 10 : maxel ; %make a v e c t o r o f e l e v a t i o n

a n g l e s85 e l = deg2rad ( e l d eg ) ; % from d e g r e e s t o r a d i a n s [ r ad ]86 D=Re . ∗ cos ( p i /2+ e l ) + s q r t ( ( ( Re+H) . ^ 2 ) −(Re . ^ 2 ) +(Re . ^2∗




90





96 C_N0=Pr−Ts−k_dB ;% C a r r i e r t o n o i s e d e n s i t y r a t i o [dB−Hz ]

97

98 %T h r e s h o l d s f o r d i f f e r e n t modu l a t i o n s as g iven i nt e c h r ep . ITU−R M.2435−0

99 BPSK_CDMA_C_N0= 3 4 . 2 ; %[ dBHz ]100 pi4_QPSK_C_N0 = 4 2 . 9 ; %[ dBHz ]101 PSK_C_N0 = 5 0 . 3 ; %[ dBHz ]102

103

104 %P l o t t i n g105 f i g u r e ( 1 )106 s u b p l o t ( 2 , 1 , 2 )107 p l o t ( r ad2deg ( e l ) ,C_N0 , ’−−r ’ , ’ LineWidth ’ , 2 )108 y l a b e l ( ’C / N_0 [ dBHz ] ’ )109 x l a b e l ( ’ E l e v a t i o n ang l e [ \ c i r c ] ’ )110 g r i d on111 t i t l e ( ’ Downlink : c a r r i e r t o n o i s e d e n s i t y r a t i o vs .

e l v a t i o n ang l e ’ )112 ho ld on113 p l o t ( r ad2deg ( e l ) ,BPSK_CDMA_C_N0. ∗ ones ( s i z e ( e l ) ) , ’ c ’

, ’ LineWidth ’ , 2 )114 p l o t ( r ad2deg ( e l ) , pi4_QPSK_C_N0 . ∗ ones ( s i z e ( e l ) ) , ’ b ’ ,

’ LineWidth ’ , 2 )115 p l o t ( r ad2deg ( e l ) ,PSK_C_N0 . ∗ ones ( s i z e ( e l ) ) , ’m’ , ’

LineWidth ’ , 2 )116 g r i d on117 l e g end ( ’C / N_0 ’ , ’BPSK /CDMA’ , ’ \ p i / 4 QPSK ’ , ’ 8PSK ’ , ’

Loc a t i o n ’ , ’ No r t hEa s t ’ )

1 %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%


2 % C a l c u l a t i o n program to e s t i m a t e Data budge t3


7 % Assumpt ions :%8 %The r o t a t i o n o f Ea r t h i s n e g l e c t e d9 % The v e l o c i t y o f t h e s h i p i s n e g l e c t e d

10 % Maximum d a t a r a t e ( raw d a t a r a t e ) i s c o n s i d e r e das c o l l e c t e d from ITU−R’ s r e p o r t M.2092−0

11 % Only one s h i p i s c o n s i d e r e d bee i ng i n t h e f i e l do f view

12 % Code r a t e i s no t t a k en i n t o a c coun t13


18 % C o n s t a n t s19 Re=6378∗10^3; % Mean r a d i u s o f Ea r t h [m]20 mu=398600 .4418∗10^9 ; % G r a v i t a t i o n a l p a r am e t e r [km

^3 / s ^2 ]21 H=550 e3 ; % O r b i t a l t i t u d e [m]22 B_R=240∗10^3; % B i t r a t e [ b / s ] , r e t i r e v e d from ITU−

R’ s r e p o r t M.2092−023




29 %Ea r t h c e n t r a l a ng l e and t ime of v i e w a b i l i t y o fs u b j e c t

30 lambda= a s i n (D . / ( H+Re ) . ∗ s i n ( p i /2+ e l ) ) ; % Ea r t hc e n t r a l a ng l e [ r ad ]

31 V= s q r t (mu . / ( H+Re ) ) ; % O r b i t a l V e l o c i t y [km / s ]32 P = (2 . ∗ p i . ∗ s q r t ( (H+Re ) . ^ 3 . / mu) ) ; % O r b i t a l p e r i o d [ s


]33 Swath_width =2 .∗ lambda . ∗ Re ; % Swath wid th [m]34 T_c =2.∗P . / ( ( 2 . ∗ p i ) . ∗ ( Re+H) ) . ∗ Swath_width . / 2 ; %Time

of v i e w a b i l i t y o f s u b j e c t [ s ]35

36 D_A=B_R∗T_c ; % The maximum ammount o f d a t a t h a t canbe t r a n s f e r e d d u r i n g one pa s s age [ b ] .

37 data_mb=D_A∗10^−6; % t r a n s f e r t o [Mb]38

39 f i g u r e ( 1 )40 s u b p l o t ( 2 , 1 , 1 )41 p l o t ( r ad2deg ( e l ) , T_c / 6 0 , ’−−b ’ , ’ LineWidth ’ , 2 )42 t i t l e ( ’ Time of c o n t a c t vs . minimum e l e v a t i o n ang l e ’

)43 y l a b e l ( ’ Time of p a s s ag e [ min ] ’ )44 x l a b e l ( ’Minimum e l e v a t i o n ang l e [ \ c i r c ] ’ )45 g r i d on46 s u b p l o t ( 2 , 1 , 2 )47 p l o t ( r ad2deg ( e l ) , data_mb , ’−−r ’ , ’ LineWidth ’ , 2 )48 t i t l e ( ’ Data ammount vs . minimum e l e v a t i o n ang l e ’ )49 y l a b e l ( ’ d a t a ammount [Mb] ’ )50 x l a b e l ( ’Minimum e l e v a t i o n ang l e [ \ c i r c ] ’ )51 g r i d on

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