Preliminary Design of a Global and Continuous Coverage ... · 4.3 Link Budget ... Iridium service...

CRANSEDS - CRANFIELD UNIVERSITY

UKSEDS-SSPI 2016 Satellite Design Competition

Preliminary Design of a Global and Continuous

Coverage Communication Services Constellation

June 2017

CranSEDS

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ABSTRACT

The following document contains a precise description of the developed

preliminary design of a constellation of small satellites, and their composing

satellites, which goal is to deliver global and continuous coverage.

This document is submitted to UKSEDS and was developed by CranSEDS team,

representing Cranfield University. The document was generated in response of

and to participate in SSPI Satellite Competition.

Keywords:

CranSEDS, Constellation, Polar Orbits, Global Coverage, Continuous Coverage,

Satellite Communications, Ka Band, GEO Interferences, Phased Array Antenna

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TABLE OF CONTENTS

ABSTRACT ......................................................................................................... i

LIST OF FIGURES ............................................................................................. vi

LIST OF TABLES ............................................................................................. viii

LIST OF ABBREVIATIONS ................................................................................ x

1 Introduction ...................................................................................................... 1

1.1 Previous Missions Baselines .................................................................... 1

1.1.1 Iridium ................................................................................................ 1

1.1.2 Globalstar ........................................................................................... 3

1.1.3 O3b .................................................................................................... 5

1.1.4 OneWeb ............................................................................................. 7

1.2 Mission Statement .................................................................................... 9

1.3 Mission Baseline ..................................................................................... 10

1.4 Mission Requirements ............................................................................ 12

1.5 Budgets ................................................................................................... 13

1.5.1 Initial Estimations ............................................................................. 14

1.5.2 Final Budgets and Comparison ........................................................ 15

2 Cost Analysis ................................................................................................. 17

2.1 Estimating Mission Lifetime .................................................................... 17

2.2 Final Cost Estimation .............................................................................. 20

3 Satellite Constellation .................................................................................... 23

3.1 Constellation Requirements and Constraints .......................................... 24

3.2 Initial Concepts & Considerations ........................................................... 24

3.3 Design of Final Constellation .................................................................. 29

3.4 Requirements and Constraints on Constellation, Including

Management of Interference ......................................................................... 32

3.4.1 Interference Mitigation ...................................................................... 32

4 Communication Subsystem ........................................................................... 35

4.1 Introduction ............................................................................................. 35

4.2 Communications Links Architecture ........................................................ 35

4.2.1 Inter-Satellite Links ........................................................................... 37

4.2.2 Gateway Stations ............................................................................. 37

4.2.3 Communication Payload .................................................................. 37

4.3 Link Budget ............................................................................................. 38

4.3.1 Frequency ........................................................................................ 40

4.3.2 Data Rate ......................................................................................... 40

4.3.3 Orbit ................................................................................................. 40

5 Regulation Aspect and Interferences Mitigation ............................................ 43

5.1 Landing Rights and Spectrum Management ........................................... 43

6 Propulsion Subsystem ................................................................................... 45

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6.1 Engine Selection ..................................................................................... 45

6.1.1 Options ............................................................................................. 45

6.1.2 Airbus 200 N Bipropellant Thruster .................................................. 48

6.2 Fuel Storage ........................................................................................... 49

6.3 Propulsion Subsystem Summary ............................................................ 50

7 Launch and Orbit ........................................................................................... 53

7.1 Launcher Selection ................................................................................. 53

7.1.1 Selection Method ............................................................................. 53

7.1.2 Atlas V 500 Series Parameters ........................................................ 55

7.1.3 Launch Procedures .......................................................................... 56

7.2 Insertion into final Orbit ........................................................................... 57

7.2.1 Computation Method ........................................................................ 59

7.3 Station Keeping and Space Debris ......................................................... 60

7.4 Delta-V and Propellant Budgets – Summary .......................................... 61

8 End of Mission Considerations ...................................................................... 63

8.1 Disposal Options and Requirements....................................................... 63

8.2 Constellation Disposal ............................................................................ 65

8.2.1 Disposal Method .............................................................................. 65

8.2.2 Computation ..................................................................................... 66

9 Attitude Determination and Control System – ADCS ..................................... 71

9.1 System Overview .................................................................................... 71

9.2 ADCS Modes .......................................................................................... 71

9.2.1 Detumbling and Data Acquisition Mode ........................................... 71

9.2.2 Normal Mode .................................................................................... 71

9.2.3 Orbit Correction Mode ...................................................................... 72

9.2.4 Safe Mode ........................................................................................ 72

9.3 Design Considerations ............................................................................ 72

9.4 Hardware Selection................................................................................. 73

9.4.1 Reaction Wheels .............................................................................. 73

9.4.2 Thrusters .......................................................................................... 74

9.4.3 Star Trackers .................................................................................... 75

9.4.4 Sun Sensors ..................................................................................... 75

9.4.5 Gyroscope ........................................................................................ 75

10 Electrical Power Subsystem ........................................................................ 77

10.1 Power Requirements ............................................................................ 77

10.2 Power Budget ....................................................................................... 77

10.3 Power Generation ................................................................................. 79

10.3.1 Primary Power ................................................................................ 79

10.3.2 Secondary Power ........................................................................... 79

10.4 Power Distribution, Management and Control ...................................... 80

10.5 EPS Mass Budget ................................................................................. 81

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11 On-Board Data Handling Subsystem .......................................................... 83

11.1 Requirements ....................................................................................... 83

11.2 OBDH Design ....................................................................................... 83

11.2.1 Architecture & Hardware ................................................................ 83

11.2.2 Memory .......................................................................................... 84

11.2.3 Protection and Fault Tolerance ...................................................... 84

11.2.4 Subsystem Interfacing .................................................................... 84

12 Structure and Configuration ......................................................................... 87

12.1 Introduction ........................................................................................... 87

12.2 Subsystem Requirements ..................................................................... 87

12.3 SSTL-150 Satellite Platform .................................................................. 87

12.4 Structure ............................................................................................... 88

12.4.1 Introduction .................................................................................... 88

12.4.2 Configuration .................................................................................. 89

12.4.3 Material .......................................................................................... 92

12.5 Configuration ........................................................................................ 92

12.5.1 Introduction .................................................................................... 92

12.6 Mechanisms .......................................................................................... 94

12.6.1 Introduction .................................................................................... 94

12.6.2 Solar Array Deployment Mechanism .............................................. 95

13 Thermal Control Subsystem ........................................................................ 97

13.1 Mission Drivers for Thermal Design ...................................................... 97

13.1.1 Overall Mission Requirements ....................................................... 97

13.1.2 Thermal Requirements ................................................................... 97

13.2 Thermal Modelling ................................................................................ 99

13.3 Conclusion .......................................................................................... 105

13.3.1 Thermal Design ............................................................................ 105

13.3.2 Further Development .................................................................... 105

REFERENCES ............................................................................................... 107

APPENDICES ................................................................................................ 111

Appendix A Atlas V 500 Series Launch System ......................................... 111

Appendix B Solar Cell Datasheet ................................................................ 112

Appendix C Battery Datasheet .................................................................... 113

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LIST OF FIGURES

Figure 1-1 Iridium satellite [1] ............................................................................. 2

Figure 1-2 Globalstar coverage [2] ..................................................................... 4

Figure 1-3 Globalstar satellite [2] ........................................................................ 5

Figure 1-4 O3b satellite [3] ................................................................................. 6

Figure 1-5 OneWeb satellite [4] .......................................................................... 8

Figure 1-6 Interferences with GEO satellites ...................................................... 9

Figure 1-7 Progressive pitching used for avoiding interferences with GEO satellites ...................................................................................................... 9

Figure 1-8 CranSEDS satellite design .............................................................. 11

Figure 1-9 Final mass budget percentages ...................................................... 15

Figure 1-10 System mass comparison between initial estimation and preliminary design mass .............................................................................................. 16

Figure 2-1 Average NASA small spacecraft mission ........................................ 18

Figure 3-1 Final constellation ........................................................................... 23

Figure 3-2 Orbit plane intersections ................................................................. 25

Figure 3-3 Polar constellation ........................................................................... 26

Figure 3-4 Polar targeting ................................................................................. 27

Figure 3-5 60° Walker delta with square beams ............................................... 28

Figure 3-6 60° Walker delta with circular beams .............................................. 28

Figure 3-7 Minimum configuration with 36° RAAN spacing .............................. 31

Figure 3-8 Minimum Configuration with 18° RAAN spacing ............................. 31

Figure 4-1 Satellite communication link architecture [10] ................................. 35

Figure 4-2 Classical satellite communication system [11] ................................ 36

Figure 4-3 Intersatellite link between the planes [12] ....................................... 37

Figure 4-4 Transparent and regenerative repeaters [10] .................................. 38

Figure 4-5 Link budget picturing [13] ................................................................ 39

Figure 4-6 Spacecraft line sight geometry [14] ................................................. 40

Figure 6-1 Airbus 200 N Bipropellant thruster sketch ....................................... 49

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Figure 6-2 Propulsion subsystem design ......................................................... 51

Figure 7-1 Hohmann transfer sampling points.................................................. 58

Figure 7-2 Spatial density > 10 cm (extracted from Operational Collision Avoidance by ESA Space Debris Office presentation given by Klaus Merz on 03/11/2016) ............................................................................................... 61

Figure 8-1 IDAC protected regions (IADC-02-01, 2007) ................................... 63

Figure 8-2 Orbital decay ................................................................................... 69

Figure 10-1 Li-Ion batteries [19] ....................................................................... 80

Figure 10-2 SST Power conditioning and distribution units [20] ....................... 81

Figure 11-1 Example of OBDH interface [21] ................................................... 85

Figure 12-1 SSTL-150 ...................................................................................... 87

Figure 12-2 Basic CAD model of the structure ................................................. 89

Figure 12-3 2D views with dimensions of the structure .................................... 91

Figure 12-4 CAD model of the satellite ............................................................. 92

Figure 12-5 CAD model with systems breakdown ............................................ 93

Figure 12-6 View of the CAD model of the satellite .......................................... 94

Figure 12-7 Double fold and roll-up solar array. Image from the University of Cambridge. ................................................................................................ 95

Figure 13-1 Fluxes impacting LEO satellite .................................................... 100

Figure 13-2 Physical characteristics ............................................................... 102

Figure 13-3 Heater system ............................................................................. 104

Figure 13-4 Louver system ............................................................................. 104

Figure 13-5 Multi-layer insulation ................................................................... 105

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LIST OF TABLES

Table 1-1 Initial mass budget ........................................................................... 14

Table 1-2 Initial power budget .......................................................................... 14

Table 1-3 Final mass budget ............................................................................ 15

Table 2-1 Estimated cost of the mission ........................................................... 22

Table 3-1: Satellites and planes required at different altitudes ......................... 29

Table 4-1 Link budget calculations ................................................................... 41

Table 6-1 Thrusters considered ........................................................................ 46

Table 6-2 MRE-1.5 & 200 N efficiency comparison .......................................... 47

Table 6-3 Airbus 200 N Bipropellant thruster technical specifications .............. 48

Table 6-4 MOOG-ISP fuel tanks considered .................................................... 50

Table 7-1 Propellant mass for orbital boost ...................................................... 54

Table 7-2 Atlas V 500 series launch capabilities into 1900 x 1900 polar orbit .. 56

Table 7-3 Constellation deployment procedure ................................................ 57

Table 7-4 Delta-V and propellant mass budgets .............................................. 62

Table 8-1 Constellation disposal trade-off ........................................................ 65

Table 9-1 ADCS system requirements ............................................................. 71

Table 9-2 ADCS hardware properties – part 1 ................................................. 76

Table 9-3 ADCS hardware properties – part 2 ................................................. 76

Table 10-1 Power budget ................................................................................. 78

Table 12-1 Main characteristics of the SSTL-150 ............................................. 88

Table 12-2 External dimensions of the structure .............................................. 90

Table 12-3 Structure breakdown ...................................................................... 90

Table 12-4 Internal volume calculation ............................................................. 90

Table 12-5 Available internal dimensions ......................................................... 91

Table 12-6 Aluminium-skinned honeycomb main properties ............................ 92

Table 12-7 Total mass and surface implemented in CATIA V5 ........................ 93

Table 12-8 Inertia matrix calculated by CATIA V5 ............................................ 94

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Table 13-1 Functional requirements of the thermal control subsystem ............ 97

Table 13-2 Temperature requirements for each subsystem ............................. 98

Table 13-3 Summary of critical phases, distances and power dissipated ...... 101

Table 13-4 Material characteristics ................................................................. 101

Table 13-5 Physical characteristics ................................................................ 101

Table 13-6 Estimated average temperature ................................................... 103

Table 13-7 Satellite temperatures per mission phase .................................... 103

Table 13-8 Thermal control subsystem power and mass budgets ................. 105

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LIST OF ABBREVIATIONS

AOCS Attitude and Orbit Control Systems

CDMA Code Division Multiple Access

CranSEDS Cranfield Students for the Exploration and Development of Space

CU Cranfield University

EPS Electrical Power System

ESA European Space Agency

FAT Frequency Allocation Table

FSS Fixed Satellite Service

GEO Geostationary Orbit

GTO Geosynchronous Transfer Orbit

IDAC Inter-Agency Space Debris Coordination Committee

ITU International Telecommunication Union

KE Kinetic Energy

LEO Low Earth Orbit

MEO Medium Earth Orbit

MLI Multi-Layer Insulation

MMH MonoMethyl Hydrazine

MSS Mobile Satellite Service

NASA National Astronautics and Space Administration

OBC On-board Computer

OBDH On-Board Data Handling

PCDU Power Conditioning and Distribution Unit

PLF PayLoad Fairing

RAAN Right Ascension of the Ascending Node

SBD Short Burst Data

SSPI Society of Satellite Professionals International

TT&C Telemetry Tracking & Commend

ULA United Launch Alliance

Introduction CranSEDS

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1 Introduction

This section introduces the performed design developed by the CranSEDS team,

from Cranfield University, participating in the UKSEDS-SSPI 2016 Satellite

Design Competition.

In this first subsection section is explained current and past projects based on

constellations of communication satellites, the mission baseline, requirements of

the mission, and budgets’ discussion. Following subsections will explain the

mission statement, baseline and requirements, concluding with the initial and final

budgets.

1.1 Previous Missions Baselines

Here is presented a group of four space missions performing global

communication activities with constellations. They are presented in chronological

order.

1.1.1 Iridium

Main Mission Parameters

Spacecraft mass 690 kg

Spacecraft power 400 W

Spacecraft lifetime 5 years

Global coverage Yes

Continuous coverage No

Number of satellites 66 (+6)

Orbit altitude 780 km

Iridium mission is operated by a USA company. This is the greatest current

constellation and continue under evolution. For example, their new generation

constellation Iridium NEXT is expected to be delivered by 2018 [1].

Its constellation is formed by 66 operational satellites, plus 6 more on-orbit

spares, distributed in 6 polar orbit planes. The designed orbit altitude is 780 km,

with an inclination angle of 86.4 degrees and 8.2 degrees of minimum elevation

angle [1]. This constellation provides global but not continuous coverage.


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Iridium service is based on user terminals communication systems. Satellites

communicate with them using an L band at 1621.35-1626.5 MHz, enabling a

telephony and modem data rate of 2.4 kbps. Each satellite has 4 inter-satellite

links operating in Ka band at 23.18-23.38 GHz and communicate with Ground

stations using Ka band links at 29.1-29.3 GHz (uplink) and 19.4-19.6 GHz

(downlink). These ground stations are called Iridium Gateways and connect the

Iridium network with the ground stablished telephone network [1].

Figure 1-1 Iridium satellite [1]

User terminals are Iridium's short burst data (SBD) transceivers that transmit data

at L band with satellites and communicate with a range of Iridium mobile phones

[1]. They provide voice and data service coverage. SBD transceivers are 400 g

weight with an antenna length of 15 cm and mean and peak transmission power

values of 0.6 and 7 W respectively.

The following information of each spacecraft subsystem is obtained from the main

webpage of the mission [1]:

• The communication subsystem uses four gimbaled nadir-pointed Ka band

antennas to transmit and receive to and from gateways. Two gimbaled Ka


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band antennas are used for East-West communications, in other words,

to communicate between orbital planes, and another two fixed Ka band

antennas are used to communicate across the same orbit plane (North-

South). Finally, 3 deployable phased antennas are used for the L band

user link.

• The power subsystems use GaAs solar arrays of 3.9 m2 and NiH2

batteries.

• Iridium’s satellite is 3 axes stabilised using momentum wheels,

magnetorquers, and thrusters. To close the control loop, they use line

horizon sensors, three axis gyros and magnetometers.

• Orbit manoeuvres are performed with a redundant system of

monopropellant hydrazine propulsion systems: a single electro-thermal

hydrazine thruster and seven hydrazine reaction engines assemblies. The

nominal fuel load of the system is 114.8 kg.

• Thermal control system is passive with electronically controlled radiators.

They use blankets and radiators.

• Finally, the structure system is based in a graphite epoxy triangular

monocoque and truss structure. Solar array deployment systems are non-

explosive, however, launch separation mechanism is pyro actuated.

1.1.2 Globalstar





Global coverage No polar coverage


Number of satellites 32

Orbit altitude 1,400 km

This project is compound of 24 second generation satellites and 8 first generation

satellites. These satellites are distributed along 8 orbital planes in 52 degrees

inclined orbits [2]. Thus, this project does not provide global coverage since poles

are not covered. Operational orbit altitude is 1400 km.


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Coverage is focused on North America, Europe, Japan and Australia. It can be

seen in Figure 1-2.

Globalstar service is provided acting as a backup coverage to users that do not

have ground stations access. A Globalstar’s satellite receive the user signal,

which can be voice or data, and is sent to the nearest ground station by the same

satellite. Then, ground network sends the information to the final user. This

Globalstar service can also be used if a ground station signal is lost to jump over

this gap in the network [2].

Figure 1-2 Globalstar coverage [2]

Communication is performed in S band achieving 1 Mbps downlink and 256 kbps

uplink for each user in second generation satellites. In total, up to 1248 different

users can be covered by each satellite. Each satellite deliver 16 different

channels divided in 78 divisions using CDMA [2].



• The communication subsystem uses S bands for service or user links,

achieving a total data rate of 1.2 Gbps. C band is used for TT&C.


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• The power subsystems use tracking solar panels and batteries for eclipse

and peak periods.

• AOCS implemented is based on 3 axis stabilisation. Utilised actuators are

momentum wheels, magnetometers and GPS. To close the control loop is

utilised sun and earth sensors.

• Orbit manoeuvres are performed with a monopropellant hydrazine based

propulsion system. The nominal fuel load of the system is 76.5 kg.

Figure 1-3 Globalstar satellite [2]

1.1.3 O3b


Spacecraft mass 700 kg (450 kg dry)

Spacecraft power 1000 W (EOL)


Global coverage No




O3b mission goal is to deliver satellite internet services and mobile backhaul

services to emerging markets.


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12 satellites form a constellation fully scalable to meet market demands. They

are found in equatorial orbits with approximately zero degrees of inclination

providing a standard coverage around +/- 45 degrees latitude. The orbit height is

8,062 km, MEO, and the mission lifetime is 10 years.

Satellites are using Ka-Band payload designed to enable the high-speed flow of

data between locations on the ground. Twelve fully steerable antennas ensure

an optimised connection to the area where data is needed. Two of them are used

to connect with gateways and the other ten to connect with users. The whole

constellation delivers 70 remote beams, each of which has a coverage of 700 km.

Each beam is capable of delivering up to 1.6Gbps of data.

Figure 1-4 O3b satellite [3]



• The communication subsystem uses 12 Ka bands steerable antennas.

• The power subsystems use two deployable three-segmented Gallium

Arsenide solar arrays and a Lithium ion battery for storage. Solar arrays

generate 1,700 W BOL and 1,000 W EOL.

• AOCS implemented is based on 3 axis stabilisation, provided by a

combination reaction wheels and magnetorquers. Attitude determination


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is provided by earth and sun sensors in conjunction with an inertial

measurement unit.

• Propulsion system applies a hydrazine monopropellant system, compound

of 8 thrusters, with 141 kg of fuel.

• OBDH implement a LEON3 microprocessor and a MIL-STD-1553B Data

Bus connecting all systems to the computer.

• The structure of the satellite is a trapezoidal in shape, consisting of rigid

aluminium honeycomb panels.

1.1.4 OneWeb



Spacecraft power Unknown

Spacecraft lifetime Unknown

Global coverage Yes

Continuous coverage Yes

Number of satellites 648 to 882


OneWeb is a mission under development that is being currently designed and

that will provide global and continues coverage to deliver internet access in every

part around the globe.

A few technical information has been released from the OneWeb mission and

satellites. It is known that the constellation is wanted to be fully operable by 2027

with an estimated investment of more than 3 billion dollars.

Constellation is estimated to be composed of 648 to 882 communication satellites

in 18 polar orbits at 1,200 km height [4].

This is the first mission in LEO that will use Ka and Ku band for communications.

Thus, specific requirements are generated in this mission in order to avoid

interferences with GEO satellites using these bands. The solution is based on

slightly spinning or tilting the satellite emission direction when passing through

equatorial latitudes. This solution is called “Progressive Pitch” and is a patent of


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OneWeb. A deeper explanation of that solution can be found in OneWeb website

[4] and clarifying representations in

Figure 1-5 OneWeb satellite [4]

Services will be provided to final users by OneWeb user terminals communicating

with the satellite. Each satellite will have the opportunity to connect with 50 to 70

ground stations called gateways.

Each satellite will be 150 kg size capable of delivering 7.5 Gbps with 17.8-20.2

GHz in gateway downlink, 27.5-30 GHz in gateway uplink, 10.7-12.75 GHz in

user downlink, and 12.75-13.25 & 14-14.5 GHz in user uplink [4].


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Figure 1-6 Interferences with GEO satellites

Figure 1-7 Progressive pitching used for avoiding interferences with GEO

satellites

1.2 Mission Statement

The proposal will be charged with providing a continuous global coverage for

internet or specific telecommunication service demanded by the customer, from

a Low Earth Orbit constellation. Customers in each country will be local

telecommunication companies which in turn provide services to private users,

government agencies or emergency and military organizations, thus making them

able to compete with other satellite communications providers and expand their

business more easily. The overarching goal is to generate profit for us and our


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telecommunication company customers through the use of innovative solutions,

based on a flexible constellation of small satellites. The target fully operative date

will be 2025, in order to compete in the market with competitors such as OneWeb.

1.3 Mission Baseline





Global coverage Yes

Continuous coverage Yes



Main spacecraft mass and power values is based on initial budgets estimations,

explained in section 1.5.1; lifetime is based on cost and revenue forecast to make

the mission profitable; and orbit parameters based on constellation design.

The designed mission constellation started with a walker delta feasibility study in

order to distribute the condensed polar coverage. After evaluation, polar orbits

showed to be more efficient and are the final choice. Moreover, this is the main

choice of developing and developed global coverage missions. These solutions

qualify for global and continuous coverage in a simple way.

Satellite design process was based on get the highest data rate possible in a 150

kg satellite. Ground stations allocation on ground is not considered in this study.

Final constellation design parameters are 11 polar orbit planes with 22 satellites

per plane. It makes a total of 242 satellites. However, for reliability issues, 3 spare

spacecraft will be delivered fulfilling all the launcher weight capabilities.

Final orbit height is 2,000 km since total mass of the mission is significantly

reduced with height and this is the maximum permitted height to comply with

client requirements to remain in LEO. It will require more propellant per launch,

but the principal launching cost factor is the total mass delivered.


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Communications will use Ka band for ground station communications at 25.3

GHz, K band for inter-satellite communications at 22.5 GHz, and S band for

telemetry tracking and command at 2.4 GHz. Two antennas are used to generate

an omnidirectional communication system for TT&C, two of the K band antennas

are used for inter-satellite communication along the same orbit and the other two

for communication with satellites in adjacent orbits, and two antennas are used

for main downlink and uplink for user communications. Regarding the data rates,

Ka band links are capable of delivering 155 Mbps, K bands 50 Mbps and S bands

10 Mbps.

In this mission, since it is using Ka bands for ground communications, will also

interfere with GEO satellites. This mission implements an array phased antenna

to change the direction of the emission and avoid these interferences. In addition,

this antennas’ technology enables the satellite to focus on covering more

demanded areas on Earth, in terms of communications.

Figure 1-8 CranSEDS satellite design

Figure above shows a simple and not fully representative model of the satellite.

In the bottom is shown the phased antenna for user communications. Solar array

is deployed in the opposite direction of the main propulsion system used for orbit

housekeeping to avoid damage on the power system due to the expelled gases.


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However, its position affects the moment of inertia and centre of mass, and must

be properly studied together with thrusters’ positions. Internally, tanks, reaction

wheels and electronic components are distributed as much symmetrically as

possible to minimize perturbing torques.

The other subsystems are top-level described below:

• The power subsystems are capable of generating 321 W EOL. It applies

GaAs NeXt Triple Junction (XTJ) Prime Solar Cells and Rechargeable Li-

Ion.

• AOCS and Propulsion systems will share their propellant tanks to minimize

the total mass of the system. It imposes that both will use Mono-methyl-

hydrazine (MMH) and N2O4 as propellants. Six bipropellant thrusters will

be applied in total. Two of them will be used for orbital manoeuvres and

the other four for the AOCS system. In total, 33.4 kg of fuel will be stored

in the satellite for orbit and attitude manoeuvres ate the beginning of the

mission.

• The thermal control system will utilise passive systems, optical solar

reflectors and cover white paint.

• To conclude, the structure will be manufactured using aluminium

honeycomb of 25 mm width based on actual designs of SSTL small

satellite blueprints. Cubic shape is chosen for better exploiting launcher

available cargo volume.

1.4 Mission Requirements

Following mission requirements are based on previous mission baseline

description and specific subsystem studies.

Functional

• The small communications satellite shall be capable of delivering 50 Mbps

of data connectivity from the LEO

• The mission shall provide continuous global coverage

• The mission shall provide inter-satellite communications


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• The satellite shall be able to maintain their orbital station

• The satellite shall be able to close the communication link to small

antennas in the ground

• Lifetime shall be estimated so that the mission is profitable

• The concept shall present at least an innovative concept or solution

• Reliability. TRL shall be not less than 7.

• The satellite shall be flexible enough to cope with different customer

needs.

• Satellite lifetime will be 8 years based on cost estimations.

Operational

• The mission shall provide a method of safely disposing the satellite at the

end of the mission life

• Satellites conforming the constellation shall have access to a ground

station network for telemetry and command purposes

• Mission design shall maximise as much as possible service profit

Constraints

• The weight of the satellite shall not exceed 150 kilograms

• The satellites shall not interfere with satellites in the GEO

• Constellation and global coverage shall be available by 2025

OneWeb is our main competitor. They plan to have the constellation and service

available by 2027.

Drivers

• The weight of the satellite shall not exceed 150 kilograms


1.5 Budgets

In this section is justified the initial estimations and explained the main changes

compared to the final design.


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1.5.1 Initial Estimations

These values are based on Wertz, et. al. [5] mass percentages breakdowns

considering that the total mass with 15% margin is 150 kg.

Table 1-1 Initial mass budget

Subsystem Percentage (%) Total Mass (kg) Margin (%)

Final Mass (kg)

Payload/Communications 31 40.4 15% 46.5

Structure 27 35.2 15% 40.5

Thermal Control 2 2.6 15% 3.0

Power 21 27.4 15% 31.5

TT&C 2 2.6 15% 3.0

OBDH 5 6.5 15% 7.5

AOCS 6 7.8 15% 9.0

Propulsion 3 3.9 15% 4.5

Other (balance + launch) 3 3.9 15% 4.5

Total 100 130.4 15% 150.0

To calculate the initial power budget, the following required power breakdown

was obtained from Jin, et. al. [6], who present a power budget for a small

communication’s satellite. Required total power was scaled using a linear relation

between mass and power.

Table 1-2 Initial power budget

Subsystem Percentage (%) Total Power (W) Margin (%)

Final Power

(W)

Payload/Communications 46 117.3 15% 134.9

Structure 1 2.6 15% 2.9

Thermal Control 10 25.5 15% 29.3

Power (including harness) 9 23.0 15% 26.4

TT&C 12 30.6 15% 35.2

OBDH 12 30.6 15% 35.2

AOCS 10 25.5 15% 29.3

Total 100 255.0 15% 293.3


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1.5.2 Final Budgets and Comparison

Final mass budget is presented below. This table was generated after the

preliminary design of each subsystem. Subsystems breakdown has been

reorganised: Launcher systems’ mass has been introduced in structure

subsystems.

Table 1-3 Final mass budget

Subsystem Percentage (%) Total Mass (kg) Margin (%)

Final Mass (kg)

Payload/Communications 8% 10.0 13% 11.3

Structure 11% 14.5 13% 16.4

Thermal Control 1% 1.0 13% 1.1

Power 17% 22.3 13% 25.3

TT&C 1% 1.0 13% 1.1

OBDH 4% 5.0 13% 5.7

AOCS 15% 20.0 13% 22.6

Propulsion 44% 58.8 13% 66.5

Total 100% 132.7 13% 149.9

Figure 1-9 Final mass budget percentages

In order to qualify for mission mass requirements, the margin available in the

system after the first design is reduced from 15% to 13%. Below are presented

the main changes in subsystem masses.


16

Figure 1-10 System mass comparison between initial estimation and preliminary

design mass

As can be seen in the graph above, propulsion and AOCS where the main

subsystems that present greater differences respect to the initial estimations.

This is mainly because the constellation was chosen to be placed in a very high

orbit with their associated transfer orbits at BOL and EOL, and the avoidance

manoeuvres, since operational orbit is at 2,000 km height.

Structure subsystem also presented a big difference respect to the initial

estimation, but in that case because the utilised materials are lighter than in

former missions.

Finally, payload and/or communications presented a reduced mass budget since

the available mass was significantly reduced due to the required propulsion

capabilities.

Cost Analysis CranSEDS

17

2 Cost Analysis

In order for the first draft of a space mission to be valuable, there is a need for a

fast assessment of its economic feasibility, as well as the technical readiness of

the industry that is needed to carry it out. This has to be taken into consideration

before a bill of materials is available.

This section contains cost estimations performed in this project. First one is based

on initial mass budget estimation and was used to calculate the lifetime of the

satellites. The last one is a proper cost estimation using preliminary design mass

values.

2.1 Estimating Mission Lifetime

A possibility for first estimates of the costs involved in a space mission can be

obtained by simply propagating the costs of reasonably similar, past space

missions. Let us consider the mission to be similar to that of surveillance or

meteorological satellites, which have an estimated historical cost of 50-150 k$/kg

[7].

Accounting for inflation, this would give us a specific cost range of 66-198 k$/kg

in 2017 US dollars. Taking the mission requirement of maximum 150 kg, and a

middle point of 132 k$/kg, the cost per spacecraft would be 19.8 million dollars.

Costs, however, may be decreased through the use of economies of scale and

learning factors during the manufacturing of the satellites, so let us consider a

cost of 100 k$/kg to account for the fact that the constellation involves a lot of

satellites, and thus a lot of opportunities for lessons to be learned.

In order to reach this costs figures, a high degree of standardisation will need to

be achieved in the manufacturing processes and the launch operations. The

implementation of COTS components and multiple-deployment launces are also

key to achieving the target cost reductions

The cost breakdown for a typical mission as provided by NASA is given in the

next figure:


18

Figure 2-1 Average NASA small spacecraft mission

With these figures, 800 satellites may be launched with a cost of 12 billion dollars.

In order for the mission to be profitable, the revenue must obviously exceed the

cost, but historically, demand estimates have been the biggest problem to the

success of this type of missions.

A Cost Per Function model is proposed to estimate the price to be charged to

private customers willing to use the system:

𝐶𝑃𝐹 =𝐼 (1 +

𝑘100)

𝑇

+ ∑ 𝐶𝑜𝑝𝑠,𝑖𝑇𝑖=1

∑ 𝐶𝑠365.25 ∗ 24 ∗ 60 ∗ 𝐿𝑓,𝑖𝑇𝑖=1

With 𝐼 being the total investment cost, 𝑘 the interest rate over 𝑇 years of life, 𝐶𝑜𝑝𝑠,𝑖

the operating costs for year I, 𝐶𝑠 the number of channels the system can support

simultaneously, and 𝐿𝑓,𝑖 is defined by:


19

𝐿𝑓 = min {

𝑁𝑢 ∗ 𝐴𝑢

365.25 ∗ 24 ∗ 60 ∗ 𝐶𝑠

1

With 𝑁𝑢 being the estimated number of users and 𝐴𝑢 being the average user

activity expressed in minutes per year.

With current day communications, a more precise measure of user activity may

be the amount of mobile data they use. According to Ericsson [8], the average

smartphone user in 2021 will use 8.9 GB of data per month, and it is expected to

grow. Fully providing this amount of data with our expected data rate would

amount to 1458.176 seconds of system operation per month, or 291.63 minutes

per year.

The same report also estimates a total worldwide mobile subscription number of

9 billion by 2021. Assuming deals are in place with local providers to back up their

ground-based services, a market share equivalent to 0.5% of this figure could

potentially be achieved.

Substituting in the formulae above for a system able to support 100000 channels

globally:

𝐿𝑓 = 0.24951

Or in a case with 0.1% of market share:

𝐿𝑓 = 0.05

From the cost breakdown, the equation for a break-even lifetime is:

𝐶𝑃𝐹 =𝐼 (1 +

𝑘100)

𝑇

+ 𝑇 ∗ 𝐶𝑜𝑝𝑠,𝑖

𝑇 ∗ 𝐶𝑠365.25 ∗ 24 ∗ 60 ∗ 𝐿𝑓,𝑖

Let us assume an accrued interest rate of 5% and cost operations equal to 8% of

total costs as per mission breakdown.


20

𝐶𝑃𝐹 =

1.2 ∗ 1010 [(1 +𝑘

100)𝑇

+ 0.08]

𝑇 ∗ 100000 ∗ 365.25 ∗ 24 ∗ 60 ∗ 𝐿𝑓

Though a typical modern-day tariff is close to a price of 1$ per 100MB, these

prices are likely to decrease, probably bringing them closer to 0.5$ per 100MB or

even lower by 2025, the intended launch date. If we assume again that we can

get 50% of this price after provider margins and taxes, and 100MB are equivalent

to 16 seconds of system operation, the system operation may cost 0.94$/min.

With the more optimistic estimate, the system would break even by its second

year of operation, the more pessimistic one ends up being profitable after its 7th

year.

The proposed lifetime of the mission is to be then 8 years, as it seems a

reasonable figure for profitability and technological obsolescence.

2.2 Final Cost Estimation

In order to provide a first order of magnitude estimate, the most common

procedure is to propagate historical data on previous space missions, preferably

some which are similar to the one being proposed. That is of course, if it is in fact

impossible to obtain data corresponding to the existing mission plan.

The model used was Aerospace Corporation’s Small Satellite Cost Model

(SSCM), which uses subsystem’s estimated weight as inputs. The reason to use

weight as a parameter is that it has been proven a reliable predictor of mission

cost.

An additional factor of 1.3 was added to the cost of the payload due to technology

readiness level considerations. TLR for this type of electronically steerable

antenna was estimated as 5, because the research done shows that, whilst it is

widely used for radar applications, it has not been flight proven for

communications purposes.


21

This model has been created to estimate the costs of small spacecraft up to 400

kg of weight, and it considers development and construction costs up to the first

flight unit. After obtaining these costs, we will assume that every subsequent unit

can be built for an approximately 90% of the same cost, and the development

costs are distributed among them. This is a rather conservative estimate, but this

fact is partially compensated by the 13% mass margin in each subsystem.

For the final constellation of 245 satellites, including spares, the cost per satellite

after the first protoflight unit comes close to 20 million dollars accounting for

inflation. Large improvements on this figure could still be made on this figure

during the development process, so if the mission can be profitable (as estimated

in section 2.1) with this estimate, it is likely to produce a better financial case.

Software is assumed to be the same for all satellites in the constellation, so its

cost is only attributed to the first unit too. From Space Mission Engineering: The

new SMAD [5], the software cost may be estimated as 500$ FY2010 per Source

Line of Code (SLOC). Previous reference for space communications missions

from the same book confirms that 40000 SLOC for the total system may be a

reasonable estimate. This is again a very conservative estimate, as most projects

are below this SLOC requirements according to the same book.

The lack of available data on operations costs makes it necessary to use another

tool to estimate them. The NASA small satellite mission cost breakdown (Figure

2-1) will be assumed to be a reasonable tool to define them, so the available cost

estimate will be added an additional 8% cost to account for these.

Cost estimates for landing rights could not be found, but they are expected to be

negligible compared to the overall program cost. Ground support equipment

development and construction are attributed to the first flight unit.

Table 2-1 presents a final overall mission cost around 6.3 billion dollars.

Presented values are updated to fiscal year 2017.


22

Table 2-1 Estimated cost of the mission

Concept First unit cost / k$

Subsequent unit cost / k$

Global costs / k$

Structure 1,296 1,165

Thermal control 382 343

ADCS 2,291 2,059

EPS 7,065 6,351

Propulsion 386 347

TT&C 607 546

OBDH 1,477 1,327

Payload 7,022 6,312

Integration Assembly & Testing

1,877 1,687

Software 24,670 0

Program management 3,092 0

Launch 863,000

Ground support 891

Total per S/C 51,057 20,138 4,964,661

Operations 466,213

Total 6,293,874

Satellite Constellation CranSEDS

23

3 Satellite Constellation

In order to meet the requirements, set for eventual global coverage, the optimal

constellation is the polar constellation. This enables the constellation to achieve

seamless global coverage that is required for any service provider

(communication, internet etc).

Figure 3-1 Final constellation

In order to stay within the Low Earth Orbit (LEO) limit as well as to cover as much

of the globe as possible with each satellite, the chosen orbit altitude for the

constellation is at 2000km. The active constellation has 11 orbital planes, 22

satellites per plane, 242 satellites in operation at any given moment. Each orbital

plane would also house two spare satellites each at lower altitudes, bringing the

total number of satellites in orbit up to 264.


24

3.1 Constellation Requirements and Constraints

The constellation for the mission is required to provide eventual global coverage

for providing data services. As such the completed constellation has to cover all

sections of the globe reliably at all times. At the same time given that global data

transfer and communication is also a part of the mission, one of the constraints

is that the satellites are required to be in view of each other, enabling inter-

satellite communication within the constellation. This will be further discussed in

the design section of the report.

Given the constraint of the mission to be in LEO, and the satellite mass limited to

a maximum of 150 kg (which restricts payload size and power), it can be initially

assumed that the number of satellites required to achieve global coverage will be

significantly large. This means that the full constellation is also going to take a

significant amount of time to be completed. Considering the operational lifetime

of each of the satellites, it is important to start to utilise them as soon as possible

to make the most of the time the spacecraft spends in orbit. Thus, a minimum

constellation configuration is advisable.

This minimum configuration will not be able to provide seamless coverage as

there will be gaps within the constellation. Thus, feasibility of the potential

minimum configurations should be evaluated in terms of service interruption.

The final consideration is that, though the service has to be global at the end of

the mission, the majority of the world’s population resides with ±60° latitude.

Thus, statistically speaking, most of the service users should also be situated

within this region. Potential constellation geometries were considered whilst

keeping this factor in mind.

3.2 Initial Concepts & Considerations

There are a limited number of standard constellation geometries available for

implementation within a mission such as this. As it is a global service providing

mission requiring inter-satellite communication, the satellites in each plane will

have to remain within observable distance of each other throughout its orbit. By


25

doing so, they are able to communicate to each other within the same plane.

Communication to satellites in other planes can be carried out nearer to or at the

orbital converging points as demonstrated below.

Figure 3-2 Orbit plane intersections

Considering these factors, the potential geometries are the Walker Delta

Constellation with additional polar orbit planes and the standard polar orbit

constellation.

A polar constellation is able to cover the surface area of the entire globe whilst

maintaining seamless, continuous coverage. Maintaining line of sight with

satellites in plane and adjacent plane is also easy in a polar constellation. This is

due to the fact that all the satellites in the constellation are affected identically by

the perturbations caused by the Earth’s Oblateness (J2 Perturbations). Thus, in

theory, the satellites need only be oriented as required at the beginning of the

mission. For the rest of the mission lifetime, they can remain as they are.


26

Figure 3-3 Polar constellation

A walker delta constellation at 60° is able to cover the heavily populate zones on

the earth. However, the Polar Regions are left without any coverage whatsoever.

This can be remedied with further orbit planes in the Polar region with further

satellites in place.

Alternatively, this can be addressed by re-directing the beams of the satellites at

higher latitudes towards the polar region. This is possible due to the fact that in a

walker delta, at the latitude of the chosen constellation inclination (i.e. 50° latitude

for a 50° inclination constellation), there is overlaps of multiple satellites within

that region.

For example, in Figure 3-6, only two of the planes are visible. This enables us to

observe that there are two satellites near the intersecting points of the orbit. As

only one satellite is required to maintain ground contact at that point, the other

satellite can be re-oriented to target the Polar Regions. As there are further

planes in the walker delta spread around the globe, a full coverage can be

achieved.


27

Figure 3-4 Polar targeting

However, a major issue remains with the Walker Delta, which is that with identical

number of satellites to that of the Polar constellation, there remain gaps within

the constellation, as shown in Figure 3-5 and Figure 3-6.


28

Figure 3-5 60° Walker delta with square beams

Figure 3-6 60° Walker delta with circular beams


29

This issue can be addressed by changing the beam shapes to geometrically

cover as much of the surface area as possible. As it can be seen, that by changing

from the square beam to circular improves the coverage.

3.3 Design of Final Constellation

Though other more novel constellation designs can be implemented to achieve

global coverage, the number of satellites required for global coverage at LEO is

considerably large. Thus, in order to reduce mission complications whilst

achieving continuous global coverage the final decision is to opt for the Polar

constellation. Though the Walker Delta can be utilised for this mission after further

parameters are implemented into the mission (changing beam shape, changing,

orientation etc), this adds further complications to the mission that are

unnecessary. As the mission is already quite involved, it is better to keep most of

the design simple.

In order to maximize the footprint of the beams on the ground the orbit altitude is

set at 2000km.

Table 3-1: Satellites and planes required at different altitudes

Altitude Satellites per plane

No. Planes

Total Satellites

2000 22.261079 11.13054 247.7778

1900 23.4327147 11.71636 274.5461

1800 24.7345322 12.36727 305.8985

1700 26.1895047 13.09475 342.9451

1600 27.8263488 13.91317 387.1528

1500 29.6814387 14.84072 440.4939

1400 31.8015414 15.90077 505.669

1300 34.2478139 17.12391 586.4564

1200 37.1017983 18.5509 688.2717

1100 40.4746891 20.23734 819.1002

1000 44.522158 22.26108 991.1113

900 49.4690645 24.73453 1223.594

800 55.6526975 27.82635 1548.611

700 63.6030829 31.80154 2022.676

600 74.2035967 37.1018 2753.087

500 89.044316 44.52216 3964.445

400 111.305395 55.6527 6194.445


30

Number of satellites required at each altitude can also be calculated roughly using

trigonometry (altitude and footprint size). As a slight overlap is required for

seamless coverage, the number of satellites comes down slightly the given value

in Table 3-1. Thus, at our chosen altitude, the required number of satellites comes

down 242 as opposed to 248.

The final consideration is the minimum configuration required for service

inauguration. Given the choice of polar constellation, any form of useable mission

structure can only be achieved when at least 5 planes have been achieved for

the mission. As the chosen Launch Vehicle (LV) is capable of launching 60

satellites at a time, within two launches a minimum configuration can be achieved.

However, some thought needs to be given to what the required geometry should

be for the minimum configuration. By spacing out the 5 planes as equally as

possible (whilst considering the completed geometry), service disruption can be

minimised to ~1hour at the equatorial region (where the disruption occurs for the

longest period of time). This means however that the LV will require higher Delta

V and subsequently more fuel to carry out the necessary plane changes. Leaving

very little margin for error this will increase the required number of launches and

thus time needed for the full constellation to be achieved.


31

Figure 3-7 Minimum configuration with 36° RAAN spacing

Figure 3-8 Minimum Configuration with 18° RAAN spacing

A better solution is to opt for the adjacent placement of the planes for minimum

configuration. This will increase service disruptions to 6 hours as opposed 1 hour,


32

however it’ll mean that the overall constellation can be achieved sooner due to

reduced number of launches.

Considering the total number of satellites required (including spares) and the

number of satellites the LV is able to launch at any given time, the entire

constellation can be completed in 5 launches, whilst the minimum configuration

can be achieved in two launches. As the chosen launch vehicle is the Atlas, which

launches around 8-10 times per year, the full constellation can be completed in

4-6 months and minimum configuration in as little as a month.

3.4 Requirements and Constraints on Constellation, Including

Management of Interference

3.4.1 Interference Mitigation

As the constellation is located in LEO and broadcasting information down to

ground using Ka bands, there will be interference with telecommunication satellite

using the same band but located in GEO. 2 ways of mitigating those interferences

have been planned.

The first one is coordinating the ITU (International Telecommunication Union) to

have specific frequencies allocated to the constellation (see section 5.1). By using

dedicated frequencies, interferences should be avoided. This is the goal of the

ITU and its spectrum management.

The second way of mitigating interference is to move the signal in a way where it

is not parallel anymore to the signal of GEO satellites. The selected way of doing

that is to use phase array antennas.

In a phased array antenna, the beam can electronically be steered in the wanted

direction. This method will allow to only move the beam and not the entire

antenna nor the whole spacecraft when crossing the GEO satellites’ beam [9].

These antennas are compound of small emitters that controlling their order or

process of emission can concentrate the broadcasted signal power in certain


33

covered areas or change the direction of emission. Basically, individual antennas

are barely delayed one respect each other to generate that effect.

Communication Subsystem CranSEDS

35

4 Communication Subsystem

4.1 Introduction

Every spacecraft does need in one way or the other some form of communication

system. The necessity for a communication system is because of the need for

telemetry, tracking and command (TT&C): for operations, or sending mission data

down to the ground (user), in most cases both forms are required [5]. A basic

communications system is made up of three sections: the space section which is

the S/C, the control station (TT&C) and the ground or mission operations. The

space section consists of onboard hardware such as: transmitters, receivers,

antennas, etc. The control section consists of TT&C facilities. The ground section

consists of high gain antennas for both reception and transmission [10].

4.2 Communications Links Architecture

Communications architecture is topology of communications links, it is the

structure of uplinks, downlinks, intersatellite links, TT&C links and ground station.

Figure 4-1 Satellite communication link architecture [10]


36

The communications architecture of this mission is designed in such a way that

it has the following:

1. Intersatellite links

2. Ground gateways stations

3. Operations ground stations (through TT&C link)

4. Voice communications, tracking and broadband access to the following

users:

a. Maritime

b. Land

c. Aviation

The designed is target every possible customer in the government, public safety,

commercial, military, personal and maritime. The general architecture will look

like the network in Figure 4-2.

Figure 4-2 Classical satellite communication system [11]


37

4.2.1 Inter-Satellite Links

To sustain a global coverage there is a need for intersatellite cross links between

satellite in same and adjacent planes with a data rate of 16 Mbps; making a star

network (link), these links will be RF links in K-band in the range 22.55–23.55GHz

[10]. Figure 4-3 shows the intersatellite link between the planes.

Figure 4-3 Intersatellite link between the planes [12]

4.2.2 Gateway Stations

As part of getting a global coverage there is a need for gateway stations to

connect customers with access to services. There will be two gateway stations

along each plane to ensure a continuous coverage and access.

4.2.3 Communication Payload

4.2.3.1 Repeaters

The communications payload consists of the repeaters and antennas, the

satellite repeaters are regenerative: unlike transparent repeaters they provide

improved link quality. On-board processing (modulation and demodulation).


38

Figure 4-4 Transparent and regenerative repeaters [10]

4.2.3.2 Antennas

The mission will constitute:

1. Four K-band phased array antenna antennas on each spacecraft for

intersatellite links

2. One Ka-band phased array antenna for transmission

3. One Ka-band phased array antenna for reception

4. Two S-band antennas for TT&C to provide omnidirectional illumination

4.2.3.3 Receivers and Transceivers

There is a need for receivers and transceivers to support the transmissions via

the antennas, the following are the chosen configuration:

1. Four K-band transceivers

2. One Ka-band receiver and transmitter each

3. One S-band transceiver

4.3 Link Budget

Link budget is the accountability of all losses involved in a communication link,

Figure 4-5 shows the idea clearly.


39

Figure 4-5 Link budget picturing [13]

Link budget is governed by the equation:

𝐸𝑏

𝑁0 =

𝑃𝐿𝑠𝐿𝑙𝐺𝑡𝐺𝑟

𝑘𝐵𝑇𝑠𝑅

Where:

P = transmitter power (W)

Ls = free space loss

Ll= other losses

Gt = transmitter gain

Gr = receiver gain

kB = Boltzmann constant (= 1.38 x 10-23 WK−1Hz−1)

Ts = system noise temperature (K)

R = data rate (bits per second)


40

4.3.1 Frequency

The frequency used:

K-band: 22.5 GHz

Ka-band: 25.3 GHz

S-band: 2.4 GHz

4.3.2 Data Rate

The date rate of the frequencies employed are:

K-band: 50 Mbps

Ka-band: 155 Mbps

S-band: 10 Mbps

4.3.3 Orbit

The orbit altitude is 2000 km; therefore, the maximum time of view is seen as in

Figure 4-6.

Figure 4-6 Spacecraft line sight geometry [14]

At 2000 km altitude and 10 ° elevation, the following parameters are:

Range max. path length: 4437 km


41

Maximum time in view: approx. 22 mins

Table 4-1 Link budget calculations

Parameter K-band Ka-band S-band

Frequency (GHz) 22.5 25.3 2.4

Data Rate (Mbps) 50 155 10

Altitude (km) 2000 2000 2000

Max. Path Length (km) 4437 4437 4437

Propagation Loss (dB) 167 208 109

Transmitter Power (W) 24 80 4

Regulation Aspect and Interferences Mitigation CranSEDS

43

5 Regulation Aspect and Interferences Mitigation

The global coverage of the constellation intends on providing service

internationally. This brings about concerns onto the regulation aspect of satellite

telecommunication: can we broadcast in another country than our own? What do

we have to do to be allowed to? This section aims at answering those questions.

5.1 Landing Rights and Spectrum Management

There are several types of license needed to operate a satellite

telecommunication system and provide service.

Satellite operator needs to get a space segment license to be able to operate.

Licensing is mandatory to regulate limited resources such as the frequencies

used and the orbital slot allocated (especially in GEO).

Landing rights refers to the need to ask authorisation to broadcast in a country.

While this still exists in several countries where it is necessary to ask for those

authorisations, the “open sky” policy renders those useless in most countries. The

ITU (International Telecommunication Union) is in charge of managing the radio

spectrum and obits allocation at an international level. This Union was created

out of a need to avoid interference when 2 satellites uses the same frequency for

example. With the “open sky” policy, once a satellite operator was granted a

license to use a part of the spectrum at the ITU level, there is no further need of

a license in another country [15,16].

Frequencies allocation is done by the government of the country where the

satellite operator is registered. As part of their spectrum management program in

line with the ITU, each country has a Frequency Allocation Table (FAT) that they

use to grant a license to a satellite operator [15,17]. Those tables describe for

each frequency available, the service that can be done, the region that can be

covered and other specifications. Services can be:

• fixed (FSS, fixed satellite service), from the satellite to a fixed station on

the ground, this is the service that we provide,

Regulation Aspect and Interferences Mitigation CranSEDS

44

• mobile (MSS, mobile satellite service), from the satellite to a mobile

station on the ground, i.e. satellite phones,

• broadcasting,

• space research and Earth observation,

• radiolocation, radionavigation, etc.

Regions are defined by the ITU:

• region 1 is Africa, Europe, Middle East and Russia,

• region 2 is the American continent,

• region 3 is Asia and Oceania.

The UK FAT has available frequencies in the Ku and Ka bands for FSS covering

all 3 regions.

The ground segment (here the small antennas) also requires licensing. The entire

licensing process both for the ground and space segments takes time and has a

cost. However, information on the delay, the cost and the ground license where

not readily available.

Propulsion Subsystem CranSEDS

45

6 Propulsion Subsystem

This chapter of the report describes the propulsion subsystem design considered

for the communication constellation proposed.

The initial section lists the options considered and their parameters as well as it

defines the approach taken in designing the subsystem. Followed by that is a

description of the fuel storage subsystem and the steps taken into sizing it,

defining its parameters and characteristics. A summary of the subsystem with a

graphical representation concludes this chapter of the report.

6.1 Engine Selection

6.1.1 Options

Space engines and thrusters can be characterised by many different parameters.

However, the most important ones are:

• Engine type, i.e. what method is used to provide thrust. Engine types vary

from solid and liquid propellant engines, through Ion and Hall Effect thrusts

to resistojets, arcjets, solar sails, etc.

• Specific Impulse (ISP) which is a measure of the engine efficiency. High

ISP corresponds to more efficient engines and vice versa.

• Thrust – defining the thrust an engine can provide. High thrust engines are

preferred for LEO space mission due to that quicker acceleration reduces

gravity losses and thus increasing manoeuvring efficiency.

• Engine mass and complexity.

• Power requirement – either for the operation of valves, turbo pumps and

mechanical devices or for molecular acceleration in the case of Ion

thrusters.

For the design of the communication constellation proposed, multiple engines

different in design and thrust generation method are considered. Generating a

list of different options is key for identifying the most feasible and most cost-

effective solution. Table 6-1 below, provides a list of the thrusters considered for


46

the mission designed. Engine thrust and mass, specific impulse as well as engine

type are listed for each of the options allowing comparison between the different

engines considered.

Table 6-1 Thrusters considered

Engine Manufacturer Thrus

t (N)

ISP Engine Mass

(kg) Type

s m/s

NSTAR Min NASA 0.019 1,700 16,67

7 43 Ion

NSTAR Max NASA 0.09 3,100 30,41

1 43 Ion

NEXT NASA 0.236 4,100 40,22

1 12 Ion

T5 Min Qinetiq 0.001 3,500 34,33

5 2.5 Ion

T5 Max Qinetiq 0.025 3,500 34,33

5 2.5 Ion

T6 Min Qinetiq 0.03 4,400 43,16

4 8.3 Ion

T6 Max Qinetiq 0.23 4,400 43,16

4 8.3 Ion

BHT-8000 Busek 0.449 2,210 21,68

0 25 Hall Effect

BHT-1500 Min Busek 0.068 1,615 15,84

3 6.8 Hall Effect

BHT-1500 Max Busek 0.179 1,865 18,29

5 6.8 Hall Effect

LEROS-1C Moog ISP 456 325 3,188 4 Bipropellant DST-11H Moog ISP 22 310 3,041 0.77 Bipropellant

TR-308 Northrop Grumman

471 322 3,159 4.76 Bipropellant

TR-312-100MN Northrop Grumman

503 325 3,188 6 Bipropellant

TR-312-100YN Northrop Grumman

556 330 3,237 6 Bipropellant

R-4D Aerojet 490 315 3,095 3.7 Bipropellant Model S400-12 Airbus 420 318 3,120 3.6 Bipropellant

200 N Airbus 216 270 2,649 1.9 Bipropellant

MRE-1.5 Northrop Grumman

86 228 2,237 1.1 Monopropellan

t

ARC-445 Aerojet 445 235 2,305 1.6 Monopropellan

t

MR-502A Aerojet 0.395 300 2,943 0.9 Resistojet

MR-509 Aerojet 0.254 502 4,925 5.5 Arcjet

At an early stage of mission design, it was agreed that liquid propulsion is the

most feasible option for the satellites considered as the other solutions do not fit


47

with mission requirements. For example, Ion propulsion demands high power

performance which is not achievable by a small spacecraft. In addition to that,

the low thrust characteristics would increase deployment time and would increase

burning time due to gravity losses.

Comparing liquid and mono propellant engines, the two feasible options were

identified. The first being the Airbus manufactured 200 N thruster and the second

being the MRE-1.5 designed and manufactured by Northrop Grumman. The

higher specific impulse of the bipropellant option meant that lower fuel mass

would be required while the monopropellant option would provide simplicity and

dry mass savings due to the lighter structure of the thruster and its related

components. For this reason, a study was conducted to select the more feasible

option of the two. The approach implemented included taking into account the

thrusts (2 thrusters for safety reasons) dry mass and the mass of the fuel required

for mission execution and then comparing the results. It must be noted that due

to the robust computational model created, different deorbiting profiles (end of life

burn) are considered. This is to comply with the requirement of complete

deorbiting within 25 years of mission finish as described in Chapter 8 of this

report.

Table 6-2 MRE-1.5 & 200 N efficiency comparison

Engine & Fuel Mass including 20% fuel margin

MRE-1.5 32.201 200 N 29.514

Difference in overall mass 2.687

As it can be seen from the results obtained, the 200 N bipropellant thruster

manufactured by Airbus is a more feasible option as it saves about 2.7 kg of

overall weight. For this reason, each satellite in the constellation will be equipped

with two 200 N engine ensuring efficiency and reduced risk in the case of engine

failure. The introduction of this type of engine to the spacecraft design simplifies

the overall orbiter design due to that the ADCS thrusters can be implemented to

the propulsion subsystem as ADCS thrusters use identical fuel and design. This


48

significantly reduces design and manufacturing issues as well as it is a more cost-

effective solution due to shared fuel tanks.

6.1.2 Airbus 200 N Bipropellant Thruster

Airbus’s 200 N bipropellant engine has been originally designed as an attitude

control, manoeuvring and braking thruster for ESA‘s Automated Transfer Vehicle.

However, due to its lightweight and relatively high efficiency the engine is a great

fit for the constellation developed. Another advantage of the engine is its reliability

as a result of the amount of missions flown. Table 6-3 Airbus 200 N Bipropellant

thruster technical specifications states the key technical parameters of the 200 N

bipropellant thruster while Figure 6-1 provides a sketch of the engine selected.

Table 6-3 Airbus 200 N Bipropellant thruster technical specifications

Airbus 200 N Bipropellant Thruster Technical Specifications

Nominal Thrust 216N ± 10N

Thrust Range 180N ± 15N to 270N ± 15N Specific Impulse at Nominal Point > 2650 Ns/kg (> 270 s)

Fuel MMH

Oxidizer N204

Fuel Density 0.8788 g/cm3

Oxidizer Density 1.447 g/cm3

Mixture Ratio Nominal 1.65 ± 0.035

Mixture Ratio Range 1.2 to 1.9 Flow Rate Nominal 78 g/s

Flow Rate Range 60 to 100 g/s Injector type Impingement with film cooling Chamber Pressure Nominal 8 bar Inlet Pressure Range 17 ± 7 bar Nozzle End Diameter (inner) 95 mm Throat Diameter (inner) 12 mm Nozzle Expansion Ratio (by area) 50 Chamber / Nozzle Material SiCrFe coated niobium alloy

Minimum on time 28 ms Minimum off time 28 ms Maximum burn time (single burn) 7600 s

Number of full thermal cycles 250

Valve power requirement 32 W Total mass 1.9 kg


49

Figure 6-1 Airbus 200 N Bipropellant thruster sketch

6.2 Fuel Storage

The introduction of the 200 N space thruster to the design of the orbiters

significantly simplifies the fuel storage subsystem due to the fuels used in

combustion and the mixture ratio of those fuels. Combustion of Monomethyl

hydrazine (MMH) and Dinitrogen tetroxide (N204) at a ratio of 1.65 defines that

the fuel tanks for the oxidiser and the fuel tanks for the fuel are of identical size.

Due to that, sizing and packaging of the fuel storage subsystem is significantly

optimised.

Fuel tanks selection is based on comparison of different pressurised fuel tanks,

space certified for MMH and N204. This is followed by computing their storage

capabilities and identifying the number of tanks required to carry the mission fuel

which is 29.9 kg in total. Three of the fuel tanks considered proved to be

adaptable to the spacecraft design. All those three fuel tanks are manufactured

by MOOG-ISP in the United States of America. Table 6-4 shows the computation

performed for each of the final options considered.


50

Table 6-4 MOOG-ISP fuel tanks considered

Fuel Tank Volume

(cm³) Mass per tank

(kg) Diameter

(cm) Tanks needed Tanks mass

(kg) Oxidizer Fuel

Hit to Kill 6489.28 1.497 23.52 1.983 1.979 5.99

Exploration 50799.90 7.666 44.59 0.253 0.253 15.34

Target 13765.13 4.218 29.4 0.935 0.933 8.44

As it can be seen, the Exploration fuel tank is too large for the mission design.

This would mean that extremely high volume of helium should be carried on for

pressurisation as well as the empty weight would be significantly higher in

comparison with the other two options. The other two options show to be capable

of carrying the fuel required for mission execution. Hit to Kill option would require

two fuel and two oxidizer fuel tanks (4 in total) while the Target fuel tank design

would require a fuel tank for each of the fuel and oxidiser. Also, both options

would be launched fuelled by over 90 % (98% for Hit to Kill and 93% for Target

fuel tank) meaning empty mass is reduced to minimum. However, due to the

lighter overall weight, the Hit to Kill option is identified as the most feasible option

for the orbiter’s fuel storage systems.

6.3 Propulsion Subsystem Summary

In summary, each of the satellites launched will be equipped with two 200 N

bipropellant thrusters as main engines. The engines will be fed by four fuel tanks

– two containing Monomethyl hydrazine (MMH) and two containing Dinitrogen

tetroxide (N204). Helium will be used for fuel tank pressurization as fuel is drained.

The fuel tanks will also feed the four ADCS thrusters. Figure 6-2 shows a

schematic representation of the propulsion subsystem of each of the satellites.

Pressure sensors, drain valves, regulators, filters, non-return safety valves and

check valves are introduced throughout the system to reduce risk and provide

control over the subsystem.


51

Figure 6-2 Propulsion subsystem design

The overall mass of the propulsion subsystem including fuel (plus 20% margin),

pipes, storage tanks (including helium) and main engines but excluding ADCS

thrusters is 40.7 kg which is about 27% of the spacecraft wet mass.

Launch and Orbit CranSEDS

53

7 Launch and Orbit

This chapter of the report describes the process implemented in launcher

selection as well as a brief summary of the launch events for the communication

mission designed. In addition to that, the deployment procedure of constellation

is discussed together with, delta V and propellant mass budgets.

7.1 Launcher Selection

7.1.1 Selection Method

Launch vehicles come in wide range of performance and capabilities. Many

launch vehicles are designed for a specific market, for example GTO or LEO.

Low Earth Orbit launchers have a significantly wider range due to the bigger

market and the ease of access in comparison to GTO or interplanetary launches.

The key parameter defining launcher selection is the final orbit at which the

launcher releases its payload (spacecraft). Based on the constellation

requirements, 242 satellites orbiting in eleven polar planes at a circular altitude

of 2000 km are required for the smooth operations of the service. However,

releasing the payload at the exact final orbit would be a complex procedure

requiring multiple manoeuvres and a long-time span to separate the spacecraft

in order to provide global coverage. For this reason, early on a decision was made

to launch the satellites into a lower orbit. Once released from the launcher, the

spacecraft would be raised into the final orbit one after another. This would

significantly reduce risk and the deployment time of the constellation.

Four different launch orbits are considered:

I. Polar 1800 x 1800 km.

II. Polar 1800 x 1900 km.

III. Polar 1900 x 1900 km.

IV. Polar 1800 x 2000 km.

Option lV was the first to be eliminated due to that it requires a more complex

launch profile and because it would cross the final (operational) altitude causing

risk to operational satellites. The launch profile was the reason for excluding the

second option from the list. Comparing options l and lll, it was confirmed that more


54

mass can be launched into the lower orbit, however this comes at a penalty due

to the increase in the mass of the fuel required to raise the orbit into a 2000 x

2000 km. Table 6-1provides the results obtained for propellant mass

requirements for the final two options considered.

Table 7-1 Propellant mass for orbital boost

Polar 1800 x 1800 km Polar 1900 x 1900 km δ

Propellant mass (kg) 4.6746 2.3347 2.3399

120% Propellant mass (kg) 5.6095 2.8016 2.8079

As it can be seen, fuel mass increase of about 100% is required when placing a

satellite into a lower orbit. For this reason, it was agreed that the satellites should

be released into the high orbit of 1900 x 1900 km.

Once the launch orbit is defined, multiple launcher vehicles, type capabilities and

specifications are studied in order to select the most feasible launcher. The

launchers considered are:

• Antares

• Atlas V

• Delta ll

• Delta lV

• Proton

• PSLV

• Rokot

• Soyuz

• VEGA

A list with performance capabilities to the selected orbit was created for each

of the launchers. Following that, comparison and trade-off was performed in

order to select the most feasible option. In addition to that, different launchers

and launcher combinations were assumed to reduce deployment period by as

much as possible. The computations performed showed that the Atlas V 500

series would be the most feasible launch vehicle option as it would allow full

deployment with six successive launches. Another important aspect taken into

account is the unprecedented success rate of Atlas V launchers ensuring low

risk and reduced insurance costs.


55

7.1.2 Atlas V 500 Series Parameters

Atlas V is a modern spacecraft launcher operated by the United Launch Alliance

(ULA) which is a joint venture between Lockheed Martin Corporation and the

Boeing Company. Atlas V uses a standard Atlas Booster, zero to five strap-on

solid rocket boosters (SRBs), a Centaur in either the Single-Engine Centaur

(SEC) or the Dual-Engine Centaur (DEC) configuration, and one of several

Payload Fairings (PLF) Invalid source specified.. A three-digit (XYZ) naming

convention was developed for the Atlas V 400 and 500 series to identify its

multiple configuration possibilities where X represents the diameter of the PLF, Y

represents the number of solid rocket boosters and Z stands for the number of

Centaur engines.

The Atlas V 400 series employs the flight-proven 4-m diameter payload fairing in

three discrete lengths: the Large Payload Fairing (LPF)—12.0m in length; the

Extended Payload Fairing (EPF) — 12.9m in length; and the Extra Extended

Payload Fairing (XEPF) — 13.8m in length. Similarly, the Atlas V 500 series

employs three lengths of the flight-proven 5.4-m diameter payload fairing: the 5.4-

m short PLF, 20.7m in length; the 5.4-m medium, 23.5m in length; and the 5.4m

long, 26.5m in length. The capability of configuring the launcher first stage based

on the mission parameters and variety of payload fairings, provides

unprecedented flexibility to the launch capabilities of the Atlas V launchers.

For the communication designed, Atlas V 500 series will be used as more

satellites can be launched with a single launch. This is possible due to the

relatively low wet mass of a single satellite. Also, as the parking orbit required is

polar, Vandenberg Air Force Base (VAFB) launch sites will be used due to its

northerly location providing easier access polar and sun-synchronous orbits.

Studying the different payload fairings of the 500 series, it was calculated that 60

satellites can be launched at once using the standard short 5.4m fairing. Due to

that short fairing provides maximum launch weight as no additional fairing mass

is carried and due to that this fits well with the mission parameters the standard

short fairing was selected for the constellation designed. In regards with payload


56

adapters, the Atlas V offers a wide range of adapters and separators varying in

size and mass. For this reason, an average value of 120 kg is allocated for

payload adapters. This mass is added to the payload mass launched at each

event. Once all launcher parameters are defined, launcher performance for the

selected launch orbit (1900 x 1900 km polar) are obtained for the different

configurations. Table 7-2 provides a summary of the launch capabilities obtained

from Atlas V Launch Services User’s Guide.

Table 7-2 Atlas V 500 series launch capabilities into 1900 x 1900 polar orbit

Launcher Launch Site Payload Fairing Mass capability (kg)

Atlas V 501 VBAF Short 5477






Error! Reference source not found. provides the key technical specifications of t

he Atlas V 500 series.

7.1.3 Launch Procedures

Ensuring launcher capabilities are maximised at every launch event is key for

ensuring cost efficiency. For this reason, different launch events using different

launcher specifications are simulated in order to obtain the optimal solution.

Based on the results obtained, the feasible solution for placing the constellation

into orbit including three spare satellites would be obtained by six launch events.

The first launch will be performed on an Atlas V 521, carrying 55 operational

satellites and a spare one. The next two launches will be performed by Atlas V

511 launchers and they carry 45 operational satellites plus a spare spacecraft at

each of the launch events. The final three sets of communication satellites will be

carried on board Atlas V 501 rockets and each event will place 33 operational

satellites. This schedule would eventually place 242 operational satellites into

orbit plus three spares in orbital storage. The spare mass at each of the launch

events would be taken by the structure housing the satellites into their position at


57

launch and will allow more spares to be launched into orbit if required at a further

stage of the design. The official rocket builder provided by ULA is used to provide

a cost estimation for each of the launch events.

Table 7-3 summarises the launch procedure implemented.

Table 7-3 Constellation deployment procedure

Launcher Satellites per Launch

Payload (kg)

105% plus adapters (kg)

Spare mass (kg)

Spares Launched

Cost Estimate (Million USD)

Atlas V 521 56 8400 9020 310 1 153 Atlas V 511 45 6750 7107.5 266.5 1 148 Atlas V 511 45 6750 7107.5 266.5 1 148 Atlas V 501 33 4950 5317.5 159.5 0 138 Atlas V 501 33 4950 5317.5 159.5 0 138 Atlas V 501 33 4950 5317.5 159.5 0 138

TOTAL 245 36750 39187.5 1321.5 3 863

7.2 Insertion into final Orbit

As stated in 7.1 above, the satellites will be released into 1900 x 1900 km polar

orbits meaning that each of the operational satellites will need to be boosted into

the higher orbit. Due to the small change in orbital altitude, standard Hohmann

transfers will be performed. The transfer will be conducted by a burn raising the

spacecraft apogee to 2000 km followed by a second circularisation burn at the

apogee. The manoeuvres will be performed in a sequence to allow spacecraft

spreading along the orbital plane. As the manoeuvres are not executed at exactly

the same point, a very small RAAN spreading would occur. However, as the

timespan is negligibly small, the operation of the constellation would not be

affected. Thus, no additional fuel for inclination changes is added to the satellites.

In the case small manoeuvres need to be carried out, 20% fuel margin is added

to each of the orbiters.

Figure 7-1 illustrates a sketch showing a standard Hohmann transfer as well as

the key sampling points for delta-V calculations.


58

Figure 7-1 Hohmann transfer sampling points

As stated in chapter 7.1.3 of this report, launchers will carry orbiters for more than

one of the orbital planes. Due to that satellites are released into a lower orbit in

comparison to the operational orbit, a RAAN computation was performed to

determine the time required for a natural switch between two orbital planes

without burning propellant. The results obtained show that orbiting at a lower

altitude for 57 orbits which corresponds to 5 days would result in 32.7 degrees

change in RAAN. This eventually means that, the satellites orbiting in the lower

orbit, would have aligned with the following orbital plane without burning any fuel

or making any orbital adjustments. This period is sufficient for planning ground

operations while keeping a tide schedule for constellation deployment.

At the end of the mission, a third burn will be performed to lower the perigee and

allow natural decay and deorbiting of the satellites. Unlike above, only one

manoeuvre will be performed, i.e. no circularization burn will be conducted at the

perigee. Detailed information and reasoning for this is provided in chapter 8 of

this report.

To summarise, two orbital manoeuvres would boost the satellites into their final

orbits while a final manoeuvre at the end of their operational life would place their

perigee deeper into the atmosphere resulting in natural orbital decay and satellite


59

burn up. As multiple satellites are launched at once, orbital planes will be

deployed one after another using the effect of RAAN spreading. Natural RAAN

spreading between two orbital planes would occur in about five days’ time which

providing sufficient time for planning and execution while keeping a tide

constellation deployment schedule.

7.2.1 Computation Method

The theory behind the computation method used is stated below while the results

are presented in Table 7-4. Calculations are performed for all three manoeuvres

required – orbit boos, circularization and de-orbit. Engine characteristics and the

mass changes due to burning fuel are considered for every next burn. 𝑉1 and 𝑉4

are the circular orbit velocity speed respectively for 1900 x 1900 km and 2000 x

2000 km. 𝑉2 and 𝑉3 are the velocities at perigee and apogee for the 1900 x 2000

km orbit. 𝑉5 is the perigee velocity for the orbit placing the spacecraft into an

uncontrolled re-entry (details provided in section 8 of this report). Based on that,

δV1 is the change in velocity placing the spacecraft into a transfer orbit while δV2

is the change in velocity required for circularisation at the operational altitude.

The final δV3 value is the required change in velocity to place the spacecraft into

de-orbiting mode. It must be noted that the first two manoeuvres are pro-grade

while the last one is retro-grade.

Velocities:

𝑉1 = √𝜇

𝑎 ; 𝑉4 = √

𝜇

𝑎

𝑉2 = √2𝜇 (1

𝑟1−

1

𝑟1 + 𝑟2) ; 𝑉3 = √2𝜇 (

1

𝑟2−

1

𝑟1 + 𝑟2) ; 𝑉5 = √2𝜇 (

1

𝑟2−

1

𝑟1 + 𝑟2)

δV:

𝛿𝑉1 = 𝑉2 − 𝑉1 ; 𝛿𝑉2 = 𝑉4 − 𝑉3 ; 𝛿𝑉3 = 𝑉4 − 𝑉5

Propellant mass:


60

𝑚𝑝 = 𝑚0 (1 −1

𝑒𝑥𝑝 (𝛿𝑉𝐼𝑆𝑃)

)

Acceleration:

acceleration =𝑇𝑡𝑜𝑡𝑎𝑙

𝑚0

Time to reach δV:

𝑎𝑐𝑐𝑒𝑙𝑎𝑟𝑎𝑡𝑖𝑜𝑛 =𝛿𝑉

acceleration

7.3 Station Keeping and Space Debris

Station keeping and debris avoidance are the two main reasons for orbital

manoeuvres in a constellation. Station keeping is required due to orbital

perturbations. In LEO, the main orbital perturbators are atmospheric drag, solar

activity, Lunar and Earth’s gravitational fields.

To start with, atmospheric drag can be ignored for orbits higher than 1000 km.

For this reason, it is assumed drag does not act on the constellation while in

operation. The other three perturbations can also be neglected due to the high

altitude and due to the relatively short operational period of the constellation. This

means that any influences of those perturbators can be ignored. Moreover, any

affections would act equally on all of the satellites in orbit meaning that the

constellation would not be affected significantly.

In regards with space debris, they vary in size and in spatial density. Of course,

larger debris are more dangerous due to the higher energy they carry.

Figure 7-2 show the spatial density of space debris bigger than 10 cm.


61

Figure 7-2 Spatial density > 10 cm

(extracted from Operational Collision Avoidance by ESA Space Debris Office

presentation given by Klaus Merz on 03/11/2016)

As it can be seen, the probability of colliding with space debris at an altitude of

2000 km is very low. Due to the low level of risk and due to the presence of spare

orbiters, debris avoidance manoeuvring can be excluded from the generic

operations planning. Especially due to that, very little fuel would be required for

such rare events.

In summary, no additional fuel for station keeping or space debris avoidance

manoeuvres is included to the propellant budgets of the satellites launched.

However, it must be noted that the extra propellant carried as margin can be used

for such manoeuvres if required, especially due to that such operations would be

exceptional and would require minimal fuel masses.

7.4 Delta-V and Propellant Budgets – Summary

As discussed above, three main manoeuvres will be executed by each of the

satellites launched into orbit. Two aiming to place the satellite into an operational

orbit and a final one for re-entry. Due to the high orbital altitude, no extra fuel is

added for orbital perturbations or debris avoidance. Table 7-4 Delta-V and

propellant mass budgets shows a summary of the delta-V and propellant budgets

calculated.


62

Table 7-4 Delta-V and propellant mass budgets

Sampling Point V1 V2 V3 V4 V5

Apogee (km) 1900 2000 2000 2000 2000

Perigee (km) 1900 1900 1900 2000 428.7

Semi-major axis (km) 8,278.14 8,328.14 8,328.14 8,378.14 7,592.49

Velocity (km/s) 6.9391 6.9599 6.8768 6.8976 6.5309

δV (km/s) 0.0208 0.0208 0.3666

Propellant (kg) 1.1737 1.1610 19.0934

120% Propellant (kg) 1.4084 1.3932 22.9121

Acceleration (m/s2) 2.8800 2.9027 2.9255

Time to reach δV (sec) 7.2219 7.1439 125.3149

The results obtained show that, the circularisation burn would require less fuel

and time in comparison to the first burn placing the spacecraft into an elliptical

orbit even though the change in velocity is identical. This is due to the lower mass

of the spacecraft at the second burn. Based in the results obtained, a total of 25.7

kg (including 20% margin) of propellant are required for orbital manoeuvring.

Additional 4.2 kg (including 20% margin) would be added to the orbiters’ fuel

tanks to serve the ADCS subsystem. In total, 29.9 kg of fuel including 20% margin

would be carried on board of each satellite.

End of Mission Considerations CranSEDS

63

8 End of Mission Considerations

8.1 Disposal Options and Requirements

The Inter-Agency Space Debris Coordination Committee (IDAC) is an

international governmental forum for addressing the issues related to man-made

space debris. The governmental body has been established in 1993 and consists

of multiple space agencies and organisations governing activities in Low Earth

Orbit (LEO) as well as issuing guidelines related to spacecraft/rocket stages

disposal. According to IDAC, responsible removal of orbiting objects from LEO is

the most major step towards limiting the growth of man-made space debris. Since

1993, the international committee studied different satellites/rocket stage

disposal options and their long-term implication on space debris mitigation.

Based on the results obtained and models generated, the committee concluded

that limiting the amount of time each object stays in LEO significantly reduces the

risk of space debris number growth.

Figure 8-1 IDAC protected regions (IADC-02-01, 2007)

IADC-02-01 Revision 1 from September 2007, defines three disposal options for

LEO satellites. The first and most preferred, is spacecraft de-orbiting at the end

of the mission. This is achieved by performing an end of life burn placing the

satellite’s perigee less than 50 km in altitude to avoid atmospheric skip. The


64

manoeuvre is also called controlled re-entry due to that, the execution must take

into account the region the spacecraft re-enters Earth’s atmosphere. This is

performed to reduce the danger factor of spacecraft pieces surviving re-entry and

causing damage to humans on Earth. IADC-02-01 requires re-entry to be

executed above oceans to reduce risks on population.

The second option defined by IADC-02-01 is disposal into a storage orbit, also

called graveyard orbit. This approach requires a satellite to be placed into a

Medium Earth Orbit (MEO). The document restricts placing satellites into

congested MEO orbits, such as 12 hour orbits and other orbits typically used by

communication and navigation satellites. Placing a spacecraft into a graveyard

orbit is achieved through series of burns raising both apogee and perigee of the

satellite into a safe storage region. Careful planning and execution is required to

eliminate collision risks and to ensure sufficient propellant is stored for the end of

life operations. This approach is useful for heavy, high altitude satellites.

In the case neither of the above options is feasible for mission design, IDAC-02-

01 recommends uncontrolled re-entry within 25 years after the end of the mission.

This is achieved by lowering the spacecraft perigee and allowing atmospheric

drag to decay the orbit naturally. As orbital decay is directly related to spacecraft

mass, IDAC recommends fuel tanks venting after the final manoeuvre is

executed. Another reason for this recommendation is to reduce risk of explosions

during the orbital decay which limits the risk of creating further space debris. In

addition to that, IDAC requires detailed studies and simulations to ensure decay

within the stated timeframe of 25 years or less. Natural orbital decay is widely

used for Earth orbiting satellites below 800 km, for large constellations and for

small satellites. While at first glance, uncontrolled re-entry might seem as the

simplest disposal method, it requires detailed planning and careful execution to

ensure responsible performance. Also, this approach adds operational

requirements as the satellite should be monitored throughout the whole period of

orbital decay to ensure decay rate, orbital position and providing warnings to

population upon re-entry over populated areas. Due to increased risk to


65

population, space agencies require detailed assessment of manufacturing

materials to ensure all hardware will burn up upon re-entry.

8.2 Constellation Disposal

8.2.1 Disposal Method

As discussed above, three different methods are commonly used for spacecraft

disposal in Low Earth Orbit. For this reason, a trade-off is conducted to select the

most feasible option. Three main parameters are selected for the study:

• Fuel mass in kilograms – considering the mass of the fuel required for end

of life manoeuvring. As the mass of the satellites is limited to 150 kg, extra

fuel would result in reduced capabilities of the spacecraft as well as issues

related to launch procedures. For this reason, Fuel Mass parameter is

allocated 80% of the trade-off performance.

• Re-entry Period in years – considering the period required spacecraft

disintegration. Longer periods would result in operational complexities and

higher risk of debris generation.

• Population Danger – is a factor related to the risk of space debris hitting

populated areas.

Based on the definition of the parameters, a method scoring lower would be the

preferable option for end of life operations. Table 8-1 provides the results

obtained from the trade-off performed for constellation disposal.

Table 8-1 Constellation disposal trade-off

Calculation example for uncontrolled re-entry:

Parameter &

Importance

Fuel Mass (kg)

80%

Re-Entry Period (Years)

10%

Population Danger

10%

Score

(Lower ==> Better)

Controlled Re-Entry 24.09 0.003 (3 hours) 2 19.48

Graveyard Orbit 5.00 200 1 24.10

Uncontrolled Re-Entry 20.11 25 3 18.89


66

𝑆𝑐𝑜𝑟𝑒 = 𝐹𝑢𝑒𝑙 𝑀𝑎𝑠𝑠 ∗ 𝐼𝑚𝑝𝑜𝑟𝑡𝑎𝑛𝑐𝑒 + 𝑅𝑒 − 𝐸𝑛𝑡𝑟𝑦 𝑃𝑒𝑟𝑖𝑜𝑑 ∗ 𝐼𝑚𝑝𝑜𝑟𝑡𝑎𝑛𝑐𝑒

+ 𝑃𝑜𝑝𝑢𝑙𝑎𝑡𝑖𝑜𝑛 𝐷𝑎𝑛𝑔𝑒𝑟 ∗ 𝐼𝑚𝑝𝑜𝑟𝑎𝑡𝑎𝑛𝑐𝑒

= 20.11 ∗ 0.80 + 25 ∗ 0.10 + 3 ∗ 0.10 = 18.89

In terms of fuel mass and danger to Earth’s population, it can be seen that

disposing the satellites into graveyard orbits is the most feasible option. However,

due to the size of the constellation and its inclination of 60 degrees which is

generally used by communication and navigation MEO satellites, graveyard

disposal is not feasible for the mission design.

Based on the results obtained, controlled and uncontrolled re-entry methods

score close results. However, the most feasible option is to allow atmospheric

drag to decay the satellite orbit as this would provide more flexibility with design

and launch operations.

8.2.2 Computation

Following the trade-off performed above, a computation is performed in order to

find the orbital parameters ensuring satellite re-entry complying with IDAC

requirements. The computation performed is based on that atmospheric drag

acts opposite to spacecraft motion and thus reducing its orbital energy. The

equation for acceleration due to drag can be defined as:

𝑎𝐷 = −1

2∗ 𝜌 ∗ 𝐶𝐷

𝐴

𝑚𝑉2

where 𝜌 is the atmospheric density, 𝐶𝐷 represents the dimensionless drag

coefficient of the spacecraft, 𝐴 stands for the cross-sectional area of the

spacecraft, 𝑚 is the mass of the satellite and 𝑉 represents the velocity of the

spacecraft. International Standard Atmosphere (ISA) defines the rate air density

decreases with altitude as follows:

𝜌 ≈ 𝜌0𝑒−𝛿ℎℎ0


67

where 𝜌 and 𝜌0 represent atmospheric density at any two altitudes, 𝛿ℎ is the

difference in altitude between those the altitudes considered and ℎ0 is the

atmosphere scale height, by which air density drops by approximately 1/𝑒.

ℎ0 is usually set to be a value lying in between 50 km and the Karman line

(approximately 100 km). Of course, atmospheric height is not constant

throughout the globe and throughout the year. Atmospheric height is influenced

by many parameters, the most important of which are gravitational fields of the

Earth and Solar activity. For uncontrolled orbital re-entries, a gravity standard can

be used. Similarly, mean values for the solar activity can be assumed when

satellite re-entry is lasting over the Sun cycle of 11 years.

Satellite mass is directly related to the re-entry time span of an orbiting body as

heavier objects carry more KE. Thus, heavier satellites require additional drag to

slow the spacecraft velocity meaning elongation in the decay period. Spacecraft

mass, cross sectional area and drag coefficients are the parameters defining the

so-called spacecraft ballistic coefficient used for orbital decay computations.

Ballistic coefficient of a body re-entering the atmosphere can be defined as

follows:

𝐶𝐵 =𝑚

𝐶𝐷 ∗ 𝐴

Standard circular orbits in LEO tend to degrade evenly due to the consistent

atmospheric drag. However, highly elliptical orbits used for uncontrolled re-entry

from higher altitudes tend to degrade differently. Due to that drag acts at perigee

and when the spacecraft is in lower altitudes, apogee decay is predominant, while

perigee decay is minimal. This sequence defined as aero braking with

atmospheric skip is repeated up to the point eccentricity is significantly reduced

with apogee in a denser atmosphere when the satellite eventually re-enters.

Computational models for uncontrolled re-entry are based on calculations

approximating changes in the semi-major axis, 𝑎 and eccentricity, 𝑒. The changes

of those orbital parameters due to drag during an orbit can be approximated to:


68

𝛿𝑎𝑟𝑒𝑣 = −2𝜋 ∗ 𝐶𝐷

𝐴

𝑚𝑎2 ∗ 𝜌𝑝 ∗ exp(−𝑐) ∗ [𝐼0 + 2𝑒𝐼 ∗1]

𝛿𝑒𝑟𝑒𝑣 = −2𝜋 ∗ 𝐶𝐷

𝐴

𝑚𝑎 ∗ 𝜌𝑝 ∗ exp(−𝑐) ∗ [𝐼1 + 𝑒

𝐼0 + 𝐼2

2]

where 𝜌𝑝 is the atmospheric density at the perigee of the ellipse, 𝑐 ≡ 𝑎𝑒/ℎ,

ℎ represent the atmospheric height and 𝐼𝑗 are the modified Bessel functions of

order 𝑗 and argument 𝑐.

Based on the method discussed above, computations and simulations were

performed in order to find the final orbital parameters ensuring atmospheric re-

entry within 25 years of mission final activities. The key parameters used for the

computation are:

• Initial apogee altitude – 2000 km.

• Drag coefficient – 2.2 (standard for satellites with a cubic body).

• Cross-sectional area – 2.67 m2 (spacecraft body and solar panels

multiplied by a factor of 0.75 taking into account change in attitude due to

atmospheric forces).

• Satellite mass – 128.57 kg. It must be noted that the computation

performed considers orbital manoeuvres and fuel burn rates. Initial

satellite mass is assumed to be 150 kg.

• Atmospheric scale height – 53 km (lower for safety reasons).

• Solar Activity – mean values.

• Orbital decay requirement – 25 years.

• Earth’s gravitational constant – 398600.4356 km3/s2.

• Earth equatorial radius – 6378.1366 km.

The result obtained show that placing a satellite with the above parameters into

a 2000 x 429 km orbit would result in natural decay due to drag in about

25 years. To be precise, the orbital parameters obtained show that the spacecraft

would be within the region of 30 km in altitude meaning that all spacecraft mass

should have burned up by this point. The results obtained comply with IADC-02-


69

01 Revision 1 from September 2007 and with ISO 24113:2011 published by the

International Organization for Standardization and prepared by Technical

Committee ISO/TC 20, Aircraft and space vehicles, Subcommittee SC 14, Space

systems and operations.

Figure 8-2 provides a graphical representation of the orbital decay showing

significant changes to the apogee and minimal changes to the orbital perigee, up

to the point both values approach merging point. Observing the figure and the

results computed, at 24 years and 9 months apogee and perigee align after which

orbital decay is relatively constant for both parameters.

Figure 8-2 Orbital decay

Attitude Determination and Control System – ADCS CranSEDS

71

9 Attitude Determination and Control System – ADCS

The purpose of the ADCS systems is to determine and control the attitude of the

satellite with high precession. The high pointing accuracy required by the antenna

payload makes the ADCS system an important part of the overall design. The

current section gives an overview of ADCS proposed design.

9.1 System Overview

The designed satellite is a small to medium, nadir pointing, telecommunication

satellite orbiting around earth at low earth orbit (LEO) at approximately 2000 km.

During the preliminary design review (PDR) the ADCS system requirement were

provided by the systems engineering and are summarized in the following table.

Table 9-1 ADCS system requirements

Description Value unit

Pointing Accuracy (Nadir) <0.5 Deg

Power budget 25 Watt

Mass Budget 20 Kg

9.2 ADCS Modes

9.2.1 Detumbling and Data Acquisition Mode

After separation from the launcher, a residual torque may exist and needs to be

cancelled by detumbling the spacecraft. The ADCS sensors are calibrated and

an acquisition of the Sun relative position is performed before using any

instrument that could be degraded by an unintended pointing toward the Sun.

This mode requires the use of inertial sensors and thrusters.

9.2.2 Normal Mode

This is the default mode of the ADCS. The system must maintain the pointing

accuracy of the antennas and compensate any environmental disturbances.

Moreover, the satellite will be oriented appropriately to maximize the solar energy

absorption by the solar panels.


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9.2.3 Orbit Correction Mode

Maintain the spacecraft on the right orbit is necessary and this mode deals with

allowing the use of the main propulsion system by pointing in the right direction

before any main boost. Moreover, error in manufacturing could lead to a

misalignment of thrusters and centre of mass that would create a torque on the

spacecraft's body during initial boost and need to be cancelled by the ADCS

system.

9.2.4 Safe Mode

The safe mode is enabled in case of failure of a control instrument. The ADCS

shall be able to get an attitude to meet the minimal housekeeping requirements

in terms of power, communication, safety of instruments and orbit.

9.3 Design Considerations

In order to select appropriate control and sensor components, the main factors

that constrain the overall design must be quantified. Therefore, the worst-case

environmental disturbances were calculated as well as the pointing accuracy

requirement.

Environmental disturbances

In Low-Earth orbit mission, environmental disturbances will affect the satellite’s

attitude through gravitational and magnetic forces or drag and pressure imposed

by external particles. These effects produce torques or velocity changes which

must be countered to achieve the right pointing or perform an accurate maneuver.

The main environmental disturbances usually considered are listed below and

impact the spacecraft's attitude relatively to its location and pointing direction in

space:

• Gravitational effect

• Magnetic waves

• Solar radiation

• Aerodynamic drag


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The ADCS system has been appropriately designed to compensate for the

disturbances that the satellite will experience at the reference orbit. The worst-

case conditions were considered.

9.4 Hardware Selection

The designed satellite is three-axis stabilized satellite. The main trade-off

hardware sizing parameters are the ones related to power efficiency, mass and

lifetime. The important requirements are the torque efficiency and the pointing

accuracy because of the payload requirements of the mission that highly rely on

these two performances.

The trade-off analysis resulted that regarding the actuators, reaction wheels were

selected for their high accuracy supported by thrusters to be able to perform agile

manoeuvres and desaturate the wheels when required or in case of emergency.

With regards to attitude determination sensors, the high accuracy of star trackers

makes them essential for the mission objectives coupled with gyroscopes that

provide inertial measurements. For redundancy, safe mode and when star

trackers cannot be used (because sensitive to bright stars and inefficient when

spinning), sun sensors have been selected.

The selected attitude control hardware is presented bellow along with its

characteristics

9.4.1 Reaction Wheels

• Smooth changes in torque allows very accurate pointing

• Nominal speed = 0

• Risk of saturation

A typical reaction wheels configuration is a pyramid of 4 identical wheels,

optimized configuration for redundancy (3-axis control still achievable even in

case of failure of one of the wheels), symmetry and balanced torque around any

axis.


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Four reaction wheels were selected, of maximum torque provided equal to 0.04

Nm and total angular momentum of 1.5 Nms.

9.4.2 Thrusters

• High-torque application

• No power needed but propellant

• Provide desaturation of wheels

Thrusters are the most frequently used attitude actuators because of their dual

use in adjusting orbital parameters such as controlling attitude, nutation, spin rate,

performing large and rapid slews and managing angular momentum. An

important advantage of using them as reaction controllers is the high level of

torque that can be obtained, dearly needed in certain tasks. Nevertheless,

attitude control accuracy is directly determined by the minimum thruster impulse

available and cannot meet high pointing accuracy requirements for typical

communication applications.

Usually, at least six thrusters are used to provide control on any axis. For our

mission purpose where ADCS thrusters are used only during detumbling mode,

safe mode and for reaction wheels desaturation, the use of only four thrusters

can achieve the same space manoeuvres. Four thrusters mounted symmetrically

about the spacecraft's center of mass provide control torques about all three

axes. A tetrahedral configuration of the thrusters allows both attitude control and

orbit control to be achieved using the same set of thrusters and makes

calculations easier because of equal torque arms along each axis.

Four 10N bi-propellant thrusters have been selected and the propellant required

for the mission has been estimated to 3.5 Kg plus 20% of margin finally equal to

4.2 Kg.

For attitude determination, the following sensors were selected and are

presented bellow along with their characteristics. High quality sensors are

required to perform accurate pointing in order to limit noise and measurement

errors.


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9.4.3 Star Trackers

• Very high accuracy

• Adapted to 3-axis stabilized spacecraft

• The vehicle must be stabilized before operating efficiently

9.4.4 Sun Sensors

• Detectors popular, accurate and reliable

• Require a clear field of view

• Very small and inexpensive

• Usable in low power acquisition and fault recovery modes

• Avoid other sensors to point toward bright stars

9.4.5 Gyroscope

• Inertial sensor that measures speed and angle changes

• To be used coupled with external sensors

• Used for nutation damping and attitude control while ring

• Smooth and high frequency information (using with external sensor)

High accuracy external sensors (star trackers and sun sensors) have been

selected and will work coupled with inertial sensors (gyroscopes) for maximum

performance. For redundancy, because they are sensitive instruments, 2 star

trackers are used and will not operate at the same time. Six sun sensors are

used, one located on each face of the spacecraft. Finally, one gyroscope MEMS

are used on each axis for a full rate measurement.

The following tables provide summarised information of the ADCS hardware

properties.


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Table 9-2 ADCS hardware properties – part 1

Subsystem Number Dimensions (mm)

Mass (Kg)

Power (W)

Temperature (deg C)

Reaction wheels

4 150x150x65 1.6 2 -20 +60

Thrusters 4 Nozzle d35 0.65 Propellant -20 +60

Star trackers 2 120x120x33 1 1.5 -40 +70

Baffle 2 d234x346 0.8 0 -40 +70

Processor 2 245x165x29 1.2 5.5 -40 +70

Sun sensors 6 95x107x35 0.21 0.1 -30 +60

Gyroscope 3 D365x123 0.06 0.4 -40 +80

Table 9-3 ADCS hardware properties – part 2

Subsystem Accuracy (deg) Torque available (Nm)

Reaction wheels 0.00028 0.04

Thrusters 1 10

Star trackers 0.00028 on pitch, yaw

0.0011 on roll

-

Sun sensors 1 -

Gyroscope 0.5/sec -

Electrical Power Subsystem CranSEDS

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10 Electrical Power Subsystem

This section describes the Electrical Power Subsystem (EPS) of the mission and

states the requirements as well as all deign considerations and decisions.

10.1 Power Requirements

The EPS has been designed to generate, store, distribute and control power to

all the satellites in the constellation. In order to ensure the EPS is designed to a

high standard, it has to fulfil the following requirements:

• The subsystem shall provide solar and battery power to each satellite

during the entire mission at all stages.

• The subsystem shall provide a control method for the power distribution to

other satellite subsystems.

• The subsystem shall provide the satellite with a ‘safe mode’ that must be

able to keep the payload and satellite in a functional state.

• The subsystem shall provide health and status data of power usage and

battery status to the On-Board Data Handling system.

10.2 Power Budget

The average and peak power required by one satellite is shown below in Table

10-1.


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Table 10-1 Power budget

Subsystem Average Power

(W)

Peak Power

(W)

Average Power +

10% margin (W)

Peak Power +

10% margin (W)

Payload 117.3 117.3 129.0 129.0

Structure 2.6 3.0 2.8 3.3

Thermal 15.5 23.5 17.1 25.9

Power 23.0 23.0 25.2 25.2

TT&C 30.6 30.6 33.7 33.7

OBDH 18.5 24.7 20.4 27.2

AOCS 25.5 25.5 28.1 28.1

Propulsion 0.0 30.0 0.0 33.0

Total 232.9 277.6 256.2 305.3

The table shows an average operational value of 256.2 W for the satellite. The

peak power shows the maximum power expected to be consumed by each

subsystem. The satellite is capable of generating 323.4 W BOL and 260.0 W

EOL.

The EPS provides power to the other subsystems. The Structures and

Mechanisms subsystem will need power for the movement of the antennas and

solar arrays; battery power will be used at the beginning of the mission to deploy

the solar arrays and antennas.

Although, the thermal system for the mission is passive, power is needed for the

electronically controlled heaters and temperature sensors. Thermal power will

vary for sunlit and eclipse periods (where thermal control is more critical).

Attitude, Orbit and Control subsystem (AOCS) comprises of reaction wheels,

magnetometer, torque rods, star trackers, GPS receiver, inertial measurement

unit earth and sun sensors that require electrical power. The Propulsion

subsystem will need power for the valve, turbo pumps and any other mechanical

devices. The On-Board Data Handling (OBDH) system uses less power for

housekeeping than during processing duties.

Power allocated to the EPS is for loss of power through the electrical harness,

interconnects and power consumption for the power conditioning and distributing


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unit. A 10% margin is added to the budget to account for components being

modified from a heritage design. The margin also accounts for any power losses

that could occur during the transfer of electrical power.

10.3 Power Generation

The main source of power is dependent on the operational environment of the

satellite and the lifetime of the mission. As the constellation is operating in LEO,

this automatically rules out the use of nuclear power, primary batteries and fuel

cells. The main source of power therefore will be solar photovoltaic solar arrays.

10.3.1 Primary Power

Gallium Arsenide NeXt Triple Junction Prime Solar Cells from SpectroLab have

are most likely to be used as they have a beginning-of-life (BOL) efficiency of

30.7% which significantly reduces the solar array area needed. These cells are

also capable of delivering 66,060 cycles and are extremely lightweight with a

mass of 2.06 kg/m2 [18]. A more detailed specification of the intended solar cell

can be seen in the Appendix.

10.3.2 Secondary Power

During periods of eclipse, the solar arrays will not be able to provide power to the

satellite. To compensate this, batteries will be used instead. Saft VL51ES Lithium-

Ion rechargeable cell has been chosen due to its high specific energy of 51 Ah,


80

very efficient, low weight and of course, the quality of Saft products in general.

The VES 16 cell specification can be seen in the Appendix.

Figure 10-1 Li-Ion batteries [19]

The cells will be placed in a battery module that will contain its own heaters as

well as a telemetry interface with the Power Conditioning and Distribution Unit

(PCDU) to provide status, battery voltage, current and temperature data.

10.4 Power Distribution, Management and Control

Electrical power will be transferred to a regulated 28 V satellite bus using a direct

energy transfer system. This system will ensure that extra power is dissipated at

BOL and that power is safely transferred. Power will be distributed through an

electrical harness that is insulted to keep heat dissipation to a minimum.


81

Figure 10-2 SST Power conditioning and distribution units [20]

Power will be managed with a PCDU from Surrey Satellites as shown above. The

unit will be radiation-hardened will protect the system from overcharging, over-

discharging, and overheating of the batteries and other satellite subsystems. In

the event of failure, the fault detection circuits in the system will be pinpoint where

the fault has occurred. Then the EPS will isolate components by cutting off power

supply using latching current limiters.

10.5 EPS Mass Budget

The EPS is comprised of four main components. In order to keep within the mass

budget that has been specified, an estimated mass budget for the power

subsystem has been conducted and can be seen in the table below.

Table 1.2 Power mass budget

Power Mass Budget Units (kg)

Solar Arrays 5.02

PCDU 4.40

Harness 3.00

Battery module 8.59

Total EPS mass 21.01

These estimations have been calculated based on the mission parameters. For

example, it has been calculated that approximately 8 battery cells (including


82

redundant cells) will be needed in order to provide 182 Wh that is needed during

the eclipse period. The solar array mass has been estimated from the array area

that is needed to provide around 300 W of electrical power.

On-Board Data Handling Subsystem CranSEDS

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11 On-Board Data Handling Subsystem

This chapter talks about the design of the On-Board Data Handling (OBDH)

system of the mission.

11.1 Requirements

The OBDH system for this satellite has to meet the following requirements in

order to ensure efficiency and reliability:

• The subsystem shall provide a method of monitoring satellite subsystems

by collecting systems telemetry.

• The subsystem shall securely store all payload and housekeeping data.

• The subsystem shall be capable of recovering from single event upsets.

• The subsystem shall send telemetry from the satellite subsystems and

deliver it to the communications system.

• The subsystem shall receive telecommands from the communications

system and deliver it to the satellite subsystems.

11.2 OBDH Design

The OBDH system is important as it is responsible for providing command and

control of all the satellite subsystems of the satellite platform and it also

commands the payload operations. Therefore, the architecture and hardware for

this mission have been carefully chosen.

The expected mass for the OBDH is 5.7 kg. This mass comprises of the data bus,

the on-board computer (OBC) and its modules and wiring and structure.

11.2.1 Architecture & Hardware

The satellite will have serial bus architecture as it is very reliable and is widely

used in the space industry. This architecture allows direct interfacing with the

OBDH system and the rest of the satellite subsystems.


84

The OBC will be centred around a LEON3 microprocessor with a MIL-STD-1553B

data bus connecting all systems to the OBC. This bus has been selected as it

has extensive flight heritage and is capable of delivering high data rates.

11.2.2 Memory

The OBD will have a dual-redundant memory system as well as random access

memory (RAM) storage that will be used for error detection and correction

(EDAC). The main memory storage will include 16 GB of storage for the payload

and a separate memory of up to 1 GB for housekeeping and utilities data.

11.2.3 Protection and Fault Tolerance

The altitude at which the satellites will be operating in are not subjected to intense

radiation, but it is still important that the components are well protected. This is

done using All the modules and equipment in the OBDH system and the OBC will

be radiation-hardened or radiation-tolerant.

In order to monitor the system well, the system will periodically perform a cycle

of reading, voting and repairing of memory. This is done to prevent a single event

upset from changing data. A voting system will be implemented in which an error

in one of the three data packets can be corrected through a ‘majority vote’. This

EDAC method and is done to prevent bit errors occurring on the same data.

11.2.4 Subsystem Interfacing

Communication between the ground and the satellite is very important. The

OBDH system allows data to be received and transmitted between the satellite

subsystems.

Upon uplink, the OBDH system will receives telecommands data from the ground.

The commands will be sent directly to the OBC to be decoded. Then it will be

sent through the data bus and directly to the desired subsystem interface.

Commands from the ground could include codes to control the AOCS

components and sensors. The OBDH system will uplink time-tagged and

position-tagged commands which will be stored and executed at a specific time


85

or when the satellite is in a certain position. The most important interface is

between the OBDH system and the communications receiver, to allow data to

enter the satellite.

Figure 11-1 Example of OBDH interface [21]

The communications system will downlink payload and telemetry data to the

ground. The OBDH system interfaces with the communications transmitter to

allow data to leave the satellite. The OBDH system interface with the AOCS is to

collect orbital data. Interface with the EPS to gather information from the PCDU

and from the Thermal system sensors to get temperature data.

Structure and Configuration CranSEDS

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12 Structure and Configuration

12.1 Introduction

This section will explain the configuration of the structure of the satellites with its

main properties and capabilities, and also the main mechanisms that are used

for movable and deployable components.

12.2 Subsystem Requirements

These are the requirements related to this subsystem that are derived from the

top-level requirements of the mission:

The mass of the satellites shall not exceed 150 kg.

The mass of the structure and mechanisms shall not exceed 20 kg.

12.3 SSTL-150 Satellite Platform

In order to ensure a feasible mission concept, looking for existing designs that

use proven technologies is a good way of doing it. For the structure of the

satellites, since the restriction in terms of mass are quite demanding, an existing

satellite platform from Surrey Satellite Technology will serve as starting point [22]:

Figure 12-1 SSTL-150


88

The SSTL-150 is a satellite platform of 153 kg designed for LEO missions that

offers a mass of 50 kg for payload [23].

These are some of the characteristics of this satellite:

Table 12-1 Main characteristics of the SSTL-150

Peak power (EOL) 100 W

Maximum Payload Mass 50 kg

Bus Dry Mass 103 kg

Mission Design Life 7 years

Types of orbits available LEO 400 km to 1000 km, any inclination

External Payload Volume 730 mm x 455 mm x 774 mm

Internal Payload Volume 279.5 mm x 231.5 mm x 252.5 mm

Structure Aluminium and aluminium-skinned honeycomb panels

Nominal schedule from Order

24 months to payload integration, 31 months to launch

Total price $18,315,000

Apart from those 3 kg that are exceeded, other aspects of the configuration will

need to be changed in order to meet the mission requirements, like the mission

lifetime, which has to be 8 years instead of 7.

Since the satellites of this mission are going to be communications satellites, the

mass that is saved for payload will be used for the subsystems most related to

the key drivers of the mission, which are the communications subsystem, the data

handling, and the power system. Since the required power for this mission is three

times the one provided by this platform, the solar panels will definitely need to be

greater.

12.4 Structure

12.4.1 Introduction

The design of the structure is very simple. It has an almost cubic shape made

with aluminium-skinned honeycomb panels.


89

In the mass estimation breakdown, the mass for the subsystem is of 20 kg taking

margins into account, so that value has not to be exceeded by both structure and

mechanisms.

This section will cover every aspect of the structure in terms of design, material

properties and impact protection capabilities.

12.4.2 Configuration

The configuration of the structure is very simple as it has been mentioned. The

external dimensions are exactly the same as the SSTL-150 [23].

Figure 12-2 Basic CAD model of the structure

The interior is divided into two spaces in order to separate subsystems that have

to be outside such as the communications and AOCS from others like data

handling, which in the case that are not radiation-hardened have to remain inside

the structure for thermal and radiation protection. It is possible then to dispense

with one of the panels to leave extra space for the antennas or the thrusters.

The interior panel has a whole in order to leave space for the fuel tanks to be in

the centre of mass of the satellite, so it does not move once the fuel starts to be

consumed.


90

Table 12-2 External dimensions of the structure

Dimensions (mm) 775 650 911 679

Thickness (mm) 25

Table 12-3 Structure breakdown

Panel Dimensions

(mm)

Volume

(m3) Quantity

Mass

(kg)

Top 650 x 679 0.011 1 1.80

Side A 775 x 911 0.018 3 5.75

Side B 650 x 911 0.015 2 4.83

Base 640 x 775 0.013 1 2.05

Total - - - 17.31

The dry mass of the basic structure is 14.43 kg using only aluminium-skinned

honeycomb panels with a density of 163 kg/m3 [24].

Table 12-4 Internal volume calculation

Available volume (Truncated pyramid)

Property Dimensions

(mm) Area (m2)

Volume (m3)

Top 625 x 654 0.41 -

Base 625 x 750 0.47 -

Height 886 - -

Total - - 0.98


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Figure 12-3 2D views with dimensions of the structure

The approximated internal volume is 0.98 m3, almost one cubic meter. The

following table shows the available dimensions for the subsystems:

Table 12-5 Available internal dimensions

Available Dimensions for Subsystems

(mm)

312.5 x 327 x

886

Available Dimensions for Payload (mm) 312.5 x 327 x

911


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12.4.3 Material

The structure is made by aluminium-skinned honeycomb panels, which core has

the following properties [25]:

Table 12-6 Aluminium-skinned honeycomb main properties

Thickness (Microns) 70

Ø honeycomb (mm) 3.2

Density (kg/m3) 163

Compressive stabilised strength

(MPa) 10.2

The panels are supported by Aluminium 6061-T6 bars that, together with the

interior panel, add stiffness to the structure.

12.5 Configuration

12.5.1 Introduction

In order to allow the control systems of the satellites to work properly, it is

necessary to integrate the different components of the structure in order to

estimate the centre of mass and the inertia matrix of the body in the most precise

way possible.

Figure 12-4 CAD model of the satellite


93

This model includes most of the subsystems with their final mass and dimensions

values:

Propulsion

Attitude Determination and Control System

Structure and Thermal Protection

Power System

Communications and Data Handling are not included for the inertia matrix

estimation, since the designs for these subsystems were not finished yet.

Figure 12-5 CAD model with systems breakdown

The representations of the Communications infrastructure consist in four

parabolic reflectors and one phased array but the designs are not accurate.

The Figure 12-5 highlights the most important parts for the inertia matrix. The

results that CATIA V5 provides with this model is shown in the next table:

Table 12-7 Total mass and surface implemented in CATIA V5

Mass (kg) 130.12

Surface (m2) 15.83


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Table 12-8 Inertia matrix calculated by CATIA V5

IXX

(kgm2) 8.82

IXY

(kgm2) - 0.18

IXZ

(kgm2) 1.41

IYX

(kgm2) - 0.18

IYY

(kgm2) 26.62

IYZ

(kgm2) 0.33

IZX

(kgm2) 1.41

IZY

(kgm2) 0.33

IZZ

(kgm2) 21.25

Figure 12-6 View of the CAD model of the satellite

As the Table 12-7 shows, the mass taken into account in the CAD model for

estimating the inertia matrix is only 130 kg, so 20 kg are missing from the total

mass budget. Those kilograms correspond to the Communications and On-Board

Data Handling mainly, but also to the rest of subsystems, since the

implementation has been done in a simplified way, but ensuring enough accuracy

to design a control system capable of working properly.

12.6 Mechanisms

12.6.1 Introduction

In order to allow the application of external infrastructure like the solar panels at

the same time that is achieved the maximum number of satellites launched per


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vehicle, deployment mechanisms are necessary, so the satellites occupy the

minimum volume possible inside the fairing.

12.6.2 Solar Array Deployment Mechanism

Due to the dimensions of the solar array (0.68 x 4.21 m), it is critical that they are

correctly stored during launch and deployed once the satellite is in the orbit.

The system applied will be the Fold-and-Roll-up Blanket with a Deployable Boom.

Defined by a study of the Defence Evaluation Research Agency and the

University of Cambridge, this system adapts properly to this case [26].

This mechanisms relays on a boom and the capability of the panels of being

folded:

Figure 12-7 Double fold and roll-up solar array. Image from the University of

Cambridge.

The solar array is folded over twice with the end bars, and then it is rolled over a

roller. A tubular boom serves as a deployable backbone. The proposed type by

the document from the University of Cambridge is a Rolatube composite boom in

order to reduce the size and mass of the deployment cassette [26].


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Thermal Control Subsystem CranSEDS

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13 Thermal Control Subsystem

13.1 Mission Drivers for Thermal Design

13.1.1 Overall Mission Requirements

Being able to connect anyone, from anywhere on Earth has become necessary

to ensure development of services and countries all around the world. This is why

global coverage is a major concern since many parts of the world cannot currently

access suitable data rate in order to exchange fast enough.

A constellation of telecommunication satellites can, located on the right orbits (cf.

Constellation WP), provide a worldwide global coverage of 50Mbp. This the main

goal of the CRANSED Team: designing a constellation of satellites able to meet

the following main requirements:

Table 13-1 Functional requirements of the thermal control subsystem

Functional requirements

• Shall be capable of delivering 50 Mbps of data connectivity

• Shall provide continuous global coverage

• Shall provide inter-satellite communications

• Shall be able to maintain their orbital station

• Shall be able to close the communication link to small antennas in the ground

• TRL shall be not less than 7.

• The satellite shall be flexible enough to cope with different customer needs.

• Satellite lifetime will be 8 years based on cost estimations.

• Satellite’s weight shall not exceed 150 kilograms

• Satellites shall not interfere with others in the GEO


13.1.2 Thermal Requirements

As a low-Earth constellation mission, the CRANSED satellites undergo harsh

thermal conditions (flux from Earth and Sun) and a thermal design is required to

ensure functional use of payload and subsystems.

Thermal Control Subsystem (TCS) must maintain the temperature of the

components of the spacecraft, and a temperature range in it. Its design highly


98

depends on the type of mission and hence, it is important to have an accurate

idea of planned orbits and components characteristics to achieve a good design.

Therefore, it is necessary to know the environment in which the spacecraft will

operate and the temperature requirements for the different components to design

a functional subsystem.

The constraints for the Thermal Control Subsystem are the ones derived from the

other subsystems, i.e. temperature requirements. Range of temperature can be

very different regarding the components or if it is located internally or externally.

Most of the components located outside the spacecraft have a wider range of

temperature limitations and can endure tougher space conditions. Most sensitive

components to temperature variations can be located in positions on the

spacecraft that provide thermal stability.

The following table summarizes thermal constraints related to each subsystem:

Table 13-2 Temperature requirements for each subsystem

Subsystem Minimal Temperature (°C) Maximal Temperature (°C)

Internal location

Antennas -100 100

Solar arrays -180 90

External location

AOCS -10 40

PCDU -20 50

Batteries 0 50

MMH Fuel -52 87

N2O4 -9.3 21.15

To estimate the thermal characteristics, the spacecraft will be sized under the

worst cases and hence we need to analyse these cases.

First, the thermal characteristics of every space environment, that the spacecraft

will encounter, need to be analysed.

These different environments depend of:


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• The power dissipated inside the spacecraft. This depends on the mission

phase and the instruments that are in use at this moment.

• The distance to the Sun that will be constant in our case and equal to 1

AU.

• The orbited body and hence the amount of albedo and IR flux received

from them. In a constellation case, only the Earth is considered.

• The distance to the orbited body. Again, the further the spacecraft orbits

from the orbited body the less albedo and IR flux influence. The orbit has

been designed to be located at 1000 Km from the Earth surface.

• Critical mission phases must be detailedto obtain the worst hot case and

the worst cold case. The most critical phases, related to thermal, are the

ones which present a high amount of power dissipated (hot case) or a low

amount (cold case) and they are executed near to the Sun (hot case) or

during eclipse (cold case). Thus, the hottest case will be the one with the

maximum input heat and the coldest one will be the one with the minimum.

13.2 Thermal Modelling

This section presents the steps taken to achieve the final design of the TCS. First,

a preliminary design will be determined; the worst cases of the mission will be

established along with a selection of the coating material.

Once all the requirements are known, it will be obtained the required

temperatures by balancing the heat dissipated inside the spacecraft through the

radiator against the heat absorbed by it. There are three kinds of external fluxes:

Solar flux, infrared flux and albedo. These last two are only considered when the

spacecraft is flying close to a body like in our constellation case.


100

Figure 13-1 Fluxes impacting LEO satellite

The balanced thermal equation used can be written as:

Where:

• Ja, Js and Jp are the albedo, the Solar flux and the IR flux, respectively.

• Aalbedo, Aplanetary and Asolar correspond to the areas which receives

different fluxes.

• Asurface is the area which radiate heat.

• a and e are the optical properties of the coating materials, absorptivity

and emissivity.

• s is the Stefan-Boltzmann constant: 5.67x10-8 W/m2K4

• Q: is the total heat dissipated inside the spacecraft.

The Table 13-3 presents a summary with the critical phases, their distance to the

Sun and the power dissipated.


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Table 13-3 Summary of critical phases, distances and power dissipated

Mission Phase Solar flux

Earth Albedo

Earth IR

Power Dissipated (W)

Worst Hot case: Sunlight + High payload use

1350 400 200 200

Worst cold case: Eclipse + Low payload use

0 0 200 20

Typical material parameters considered for calculation:

Table 13-4 Material characteristics

Coating Material Absorptivity Emissivity

2mil Aluminized Teflon 0.1

0.66

8mil Quartz Mirrors 0.05

0.8

1⁄2 mil Aluminized Kapton

0.34

0.55

2mil Aluminized Kapton 0.41 0.75

Honeycomb aluminium panel

0.1

0.8

Table 13-5 Physical characteristics

view factors

F (earth face) F (side faces)

0.8 0.2


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Figure 13-2 Physical characteristics

On-board temperatures are obtained using the balanced thermal equation. The

following table shows the average temperature obtained for each coating

material, in the hot (red) and cold (green) case scenarios.


103

Table 13-6 Estimated average temperature

Solar flux Js Albedo flux Ja Earth IR flux Je

Internal heat T(°K) Coat material

1361.63 400 200 200 228.43 2mil Aluminized

Teflon

1361.63 400 200 200 212.61 8mil Quartz Mirrors

1361.63 400 200 200 274.12 1⁄2 mil Aluminized

Kapton

1361.63 400 200 200 265.13 2mil Aluminized

Kapton

1361.63 400 200 200 221.59 Honeycomb

aluminium panel

0.00 0 200 20 179.13 2mil Aluminized

Teflon

0.00 0 200 20 178.45 8mil Quartz Mirrors

0.00 0 200 20 179.90 1⁄2 mil Aluminized

Kapton

0.00 0 200 20 178.67 2mil Aluminized

Kapton

0.00 0 200 20 178.75 Honeycomb

aluminium panel

For structural considerations, honeycomb aluminium panels will be used because

of high stiffness and reliability.

Finally, we obtain in °C:

Table 13-7 Satellite temperatures per mission phase

Mission Phase Temperature (°C)

Worst Hot case:

Sunlight + High payload use -52

Worst cold case:

Eclipse + Low payload use -95

We can see that for typical environment, none of these coating material is able to

keep the subsystems in the right thermal conditions. Indeed, the temperature is

always too low, even under sunlight especially with honeycomb aluminium panel


104

with very low absorptivity. This is why, it is necessary to add systems to heat up

the overall spacecraft temperature.

Some examples of active control are provided:

• Heaters: these components are used for heat up different equipment

during the coldest case along the mission and therefore avoiding huge

fluctuations in temperature. They may include some thermostat to control

the temperature of a specific component.

Figure 13-3 Heater system

• Louvers: these active thermal control components allow the rejection of

internal heat when they are open and they avoid it when they are closed.

So, they can control how much internal heat it is going to be dissipated.

Figure 13-4 Louver system

Patch heaters will be used for their simplicity (no mechanical actuators needed)

and reliability. They will use heating resistors to increase the temperature of the

spacecraft.

MLI (Multi-Layer Insulation) will be used to limit temperature losses as well as

black paint to increase absorptivity and decrease emissivity to have a relatively

good efficiency and some thermal inertia while heating up.


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Figure 13-5 Multi-layer insulation

13.3 Conclusion

13.3.1 Thermal Design

To conclude, a simple preliminary thermal design has been done using the

thermal balanced equation and the subsystem and honeycomb aluminium is

required for structure purposes. The overall temperature of the spacecraft is too

low and does not fulfil the payload requirements without the use of active thermal

systems even with MLI to limit thermal fluxes with the environment and black

paint to slightly change the spacecraft’s thermal. This is why patch heaters will

be used and consume power to heat up the overall satellite.

Typical mass and power budget for simple thermal subsystem:

Table 13-8 Thermal control subsystem power and mass budgets

Component Mass (Kg) Power (W)

MLI 3 0

Heaters 1 20

paint 1 0

13.3.2 Further Development

Thermal simulations on software to improve the design as well as testing in

vacuum and thermal cycle chamber must be done before flight.

REFERENCES CranSEDS

107

REFERENCES

1. Iridium Communications Inc. Iridium. 2017. Available at:

https://www.iridium.com/ (Accessed: 20 July 2001)

2. Globalstar. Globalstar. 2017. Available at: https://www.globalstar.com/en/

(Accessed: 1 January 2017)

3. O3b Networks. O3b Networks. 2017. Available at:

https://www.o3bnetworks.com/ (Accessed: 1 January 2017)

4. OneWeb. OneWeb. 2017.

5. Wertz JR., Everett DF., Puschell JJ. Space mission engineering : the new

SMAD. 2011.

6. Jin JIN., Linling K., Jian YAN., Xi C., Zuyao NI., Xiaochuan YOU., et al.

SSC15-V-2 Smart Communication Satellite ( SCS ) Project Overview.

2014;

7. MIT. Communications Satellite Constellations. Wireless World. 2003;

8. Ericsson. Mobility Report: on the Pulse of the Networked Society.

Automatic Documentation and Mathematical Linguistics. 2016; Available

at: DOI:10.3103/S0005105510050031

9. Perko K et al. X-Band phased array antenna validation report. 2002;

10. Maral, G., Bousquet, M., and Sun Z. Satellite Communication Systems.

5th edn. Chichester: Wiley (ed.) 2009.

11. SlidePlayer. Satellite Communication System. 2016. Available at:

http://slideplayer.com/slide/5787016/ (Accessed: 1 June 2017)

12. Yang, Qing, Yu, Wei, Challal Y. A Quasi-Dynamic Inter-Satellite Link

Reassignment Method for LEO Satellite Networks. 11th International

Conference, WASA. 2016;

13. Tsai M. Link Budget. Available at:

REFERENCES CranSEDS

108

https://www.csie.ntu.edu.tw/~hsinmu/courses/_media/wn_11fall/link_budg

et.pdf (Accessed: 1 January 2017)

14. Hoffman EJ. Space Communications. Oxford University.

15. International Telecommunication Union (ITU). Article 5 of the Radio

Regulations. 2004;

16. International Telecommunication Union (ITU). Regulation of global

broadband satellite communication. 2012;

17. National Frequency Planning Group. United Kingdom Frequency

Allocation Table. 2013;

18. Spectrolab. Space products: Cells. 2016. Available at:

http://www.spectrolab.com/DataSheets/cells/XTJ_Prime_Data_Sheet_7-

28-2016.pdf

19. Saft. MP & SMALL VL. 2017. Available at:

http://www.saftbatteries.com/battery-search/mp-small-vl

20. Surrey Satellite Technology US. Power Conditioning and Distribution Unit

Datasheet. 2016. Available at: https://www.sst-us.com/getfile/68106d20-

6030-4a51-9792-d1141306acf2

21. Addaim. A, Kherras. A and ZEB. Design of Low-cost

Telecommunications CubeSat-class Spacecraft. INTECH Open Access

Publisher. 2010;

22. NASA. SSTL-150 Satellite Platform.

23. Surrey Satellite Technology. NASA Rapid III SSTL-150 ICD File. 2011;

24. Swiss Composite. Aluminium Honeycomb. 2016; : 782505.

25. Rao KK., Rao JK., Gupta KSA. Heat Insulation Analysis of an Aluminium

Honeycomb Sandwich. Journal of Thermal Engineering. 2015; 1(3): 210–

220.

REFERENCES CranSEDS

109

26. Pellegrino S., Kukathasan S., Tibert G., Watt A. Small Satellite

Deployment Mechanisms. 2000;

APPENDICESAtlas V 500 Series Launch System CranSEDS

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APPENDICES

Appendix A Atlas V 500 Series Launch System

Figure A 1 Atlas V 500 series launch system

APPENDICESSolar Cell Datasheet CranSEDS

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Appendix B Solar Cell Datasheet

Figure A 2 SpectroLab 30.7% NeXt triple junction (XTJ) prime solar cells

datasheet [18]

APPENDICESBattery Datasheet CranSEDS

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Appendix C Battery Datasheet

Figure A 3 Saft VL51ES Li-Ion cell datasheet [19]

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