1

NANOSTAR consortium

Space Mission PredesignChallengeSurvivability in space of Roscoff worms

STAR WORMS UC3M teamCarlos Álvaro Arroyo ParejoMiguel Muñoz LorenteMiguel Renieblas AriñoÁlvaro Sanz CasadoSergio Sarasola Merino

Executive AbstractThe Nanostar challenge proposes a predesign of a scientific nanosatellite to verify the survivability in spaceof Roscoff worms with the objective of assessing their potential capability of creating artificial ecosystemsfor future deep space exploration missions. The proposed design solution has to carry a minimal scientificpayload that consists of a closed environment artificial ecosystem that is able to maintain alive, monitorthe metabolism and retrieve data of the worms colony.

The mission utilises a CubeSat that carries the biological payload which includes all the systems tomaintain alive the self-sustained colony of Roscoff worms. It deals with the quality and flow of artificial-sea water, the levels of oxygen and carbon dioxide, the quality of the light spectrum and the quantityof photons released for the photosynthesis of the "animalgae", the regulation of temperature and theregulation of the associated microbiome. For this purpose, the payload is equipped with a set of probes, apumping system, an artificial light provided by means of a 1W-LED, a full HD resolution camera andtemperature control devices.

In order to develop a solution which can accomplish this ambitious scientific mission, a concurrent designstudy is performed. Following the CubeSat philosophy, the objective is to find the most robust, maturedand reliable solution that complies with the payload needs and requirements. The iterative design processcarried out by the Star Worms UC3M team has led to an optimum solution in terms of orbit selection andsystems design and sizing.

The mission implements a carefully analysed orbit design: The S/C is introduced in a SunsynchronousLow Earth Orbit. With this, firstly, the eclipse periods are minimised which allows a more uniform thermalenvironment along the mission. Secondly, the power generation and power distribution strategies aremore robust and less complex. Thirdly, the use of a propulsion system is avoided, which generally has thelowest maturity level in CubeSats and implies an important risk for a biological payload. And last but notleast, end-of life regulation is to be accomplished naturally by interaction of the S/C with the atmosphere.

The solution of the performed concurrent engineering study is a 6U Cubesat (2Ux3U) with 16kg mass.Steady power generation is achieved with a solar array and eclipse-free orbits. An accurate pointingis achieved by means of an star tracker and reaction wheels (3-axis stabilised). The CubeSat is alsoequipped with a set of fine sun-sensors to allow a quick attitude determination after detumbling and in safemode. Reaction wheels are desaturated utilising a 3-axis magnetorquer. The payload has a dedicated highspeed on-board computer to process, compress and store HD photo/video. Communication with Earth isachieved through a reliable UHF antenna and transponder. Payload data is sent to the ground stations bymeans of an S-Band Antenna and its corresponding transceiver. A semi-active thermal control strategywith advanced coatings and heaters is considered for units’ temperature control. Finally, a lightweightstructure and radiation shielding contain all the components.

UNIVERSIDAD CARLOS III DE MADRID

January 2020

II

Change Log Record

Edition/Revision Date Description of the changeV0.0 10/10/2019 Initial version of the document

V0.1 15/11/2019Added Sections 3.4 Requirements Flowdownand 3.5 Acceptance Tests, minor changes in allsubsystems

V0.2 27/12/2019Updated description of Thermal Control Subsystemafter High-Level requirements second release

V0.3 30/12/2019 Added Appendix B. System Data Summary

V0.4 02/01/2020Added Executive Abstract and Sections 6. RiskAnalysis and Mitigation and 7. Summary andConclusions

V1.0 05/01/2020 Final version of the document, minor corrections

Table 1: Change Log Record.

III

AcknowledgementsFirstly, we want to thank NanoStar Challenge organisers and Interreg Sudoe for creating and encouragingthis competition. As a team, we have not only learnt a lot about space engineering, but also we have hadso much fun during these months. We want to specially thank Filippo Cichocki and Olivier Marty fortheir work and implication with all the students during the whole challenge.

Besides, we want to express our gratitude to UC3M professors Matilde Pilar Sánchez Fernández andEmanuele di Sotto for their support in Communications and GNC topics respectively.

IV

Contents

Executive Abstract . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II

Change Log Record . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III

Acknowledgements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IV

Contents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VII

List of Figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VIII

List of Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IX

Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . XI

Applicable documents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . XII

Reference documents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . XIV

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

1.1 Objective and Scope 1

1.2 Structure of the Document 1

1.3 Team Composition and Roles 1

2 Mission Overview, Requirements and Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2.1 Mission Overview 32.1.1 Mission Statement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.1.2 Mission goals, Payload and Predesign Ideas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.1.3 Summary of the Design Solution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.1.4 Mission Phases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.2 Mission Environment 6

2.3 Project Management 7

2.4 Requirements Flowdown 8

2.5 Acceptance Tests 8

3 Concurrent Engineering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

4 Subsystems: Analysis and Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

4.1 Mission Analysis 124.1.1 De-Orbit Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.1.2 Eclipse Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.1.3 G/S Coverage Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.2 Mission Architecture 15

4.3 Payload 16

4.4 System Operation Modes 17

4.5 Communications Subystem (CS) and Ground Segment (GS) 174.5.1 Frequency selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174.5.2 COTS trade-off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174.5.3 Link budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

4.6 Command & Data Handling 21

4.7 Electrical Power Subsystem (EPS) 214.7.1 Power Generation Strategy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224.7.2 Power Budget and Power Operational Modes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224.7.3 Solar Array and Battery Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224.7.4 COTS Trade-off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.7.5 S/C connection diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

4.8 Attitude, Determination and Control Subsystem (ADCS) 274.8.1 Disturbance sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 274.8.2 Selection and Sizing of ADCS components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 304.8.3 COTS trade-off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

4.9 Mechanical design and structure 344.9.1 Mechanical Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.9.2 S/C Shielding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

4.10 Thermal Control System (TCS) 374.10.1 Thermal Mathematical Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 374.10.2 TCS design: iteration and solution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 404.10.3 Requirements Compliance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

4.11 S/C configuration 42

4.12 System Budgets and CubeSat Deployer 44

5 Risk Analysis and Mitigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

6 Summary and Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

6.1 Future Works 49

A Mission Requirements and Compliance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

B N2chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

C COTS Datasheet Links . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

D Project Management Charts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

E Coverage and Link Budget Charts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

VI

F Thermal Control Subsystem Charts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

F.1 Model representation 63

F.2 Hot case: temperature evolution of all nodes along time 64

F.3 Cold case: temperature evolution of all nodes along time 66

G Utilised Resources . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69

H System Data Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

VII

List of Figures

1.1 Star Worms UC3M Organization Breackdown Structure. . . . . . . . . . . . . . . . . . . . 2

2.1 Top Level Operations Breakdown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

3.1 Concurrent Design Flowchart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

4.1 Re-entry simulation for a surface area of 0.02 m2. . . . . . . . . . . . . . . . . . . . . . . . 124.2 Representation of a sun synchronous orbit. . . . . . . . . . . . . . . . . . . . . . . . . . . . 134.3 Eclipse time in an orbit along the year for different altitudes. . . . . . . . . . . . . . . . . . 134.4 Orbit ground track and ESA CORE Network coverage simulation. . . . . . . . . . . . . . . 144.5 System Architecture. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.6 S/C Components Connection Diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.7 Spacecraft nominal attitude in orbit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 274.8 Temperature ranges according to ECSS-E-ST-31C. . . . . . . . . . . . . . . . . . . . . . . 374.9 Thermal Results Iterations. Top-left 1st It; Top-right 2nd It; Bottom-left 3rd It; Bottom-

right 4th It. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 404.10 S/C Stack Configuration (left) and S/C Deployed configuration (right). . . . . . . . . . . . . 434.11 S/C layout. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

B.1 N2 chart functioning diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56B.2 N2 chart. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

D.1 Project Work Break Down Structure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59D.2 Project Work Flow. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

E.1 Contacts with ESA CORE Network in 24h. . . . . . . . . . . . . . . . . . . . . . . . . . . 61E.2 Signal attenuation due to the atmosphere. . . . . . . . . . . . . . . . . . . . . . . . . . . . 61E.3 Signal attenuation due rain. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62E.4 Energy per bit to noise ratio required for the desired BER. . . . . . . . . . . . . . . . . . . 62

F.1 Representation of the radiation connections between nodes. . . . . . . . . . . . . . . . . . . 63F.2 Representation of the conduction connections between nodes . . . . . . . . . . . . . . . . . 63F.3 Interior panels temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64F.4 MLI temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64F.5 Solar panels temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64F.6 Antennas temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65F.7 Electronic devices temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65F.8 Electronic controllers temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65F.9 Battery temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66F.10 Payload temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66F.11 Solar panels temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66F.12 Antennas temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67F.13 Electronic devices temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67F.14 Electronic controllers temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67F.15 Battery temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68F.16 Payload temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

List of Tables

1 Change Log Record. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . III2 Applicable Documents. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . XII

2.1 Mission Fact-sheet. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.2 Top Level Requirements verification guidelines. . . . . . . . . . . . . . . . . . . . . . . . . 92.3 Acceptance Tests Proposal. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

4.1 Contact time per day for different orbit altitudes. . . . . . . . . . . . . . . . . . . . . . . . 144.2 Trade-off study for S-Band transceiver. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.3 Trade-off study for UHF transceiver. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.4 Trade-off study for S-Band patch antenna. . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.5 Trade-off study for UHF antenna. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184.6 S-band downlink RF link budget. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204.7 Power Budget and Power Operational Modes. . . . . . . . . . . . . . . . . . . . . . . . . . 234.8 Solar Array Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.9 Battery Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.10 Solar Array Trade-off. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.11 Battery Trade-off. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.12 PMAD Trade-off. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.13 Disturbance sizing for the ADCS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 304.14 Trade-off study for reaction wheels. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.15 Trade-off study for magnetorquer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.16 Trade-off study for star trackers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.17 Trade-off study for sun sensors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.18 Shielding estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 364.19 Clarification of the variables included in the thermal balance equation. . . . . . . . . . . . . 384.20 Comparison between Hot and Cold cases. . . . . . . . . . . . . . . . . . . . . . . . . . . . 404.21 Design parameters of the TCS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 414.22 Temperatures at each node. Compliance with requirement NS-SYS-05 . . . . . . . . . . . . 424.23 S/C Mass, CoG and Inertia without margins in Deployed config. . . . . . . . . . . . . . . . 434.24 S/C Mass, CoG and Inertia with component margins in Deployed config. . . . . . . . . . . . 434.25 System Mass Budget. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 454.26 Procurement Cost Budget. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

5.1 Risk and Mitigation analysis. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

6.1 Mission and S/C characteristics. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

A.1 Requirements list, Parent requirement and Verification. . . . . . . . . . . . . . . . . . . . . 51A.2 Requirements Flowdown Table . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

C.1 COTS Datasheet links. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

Nomenclature

ADCS Attitude Determination and Control Subsystem

BOL Beginning of Life

C&DH Command and Data Handling

CoG Centre of Gravity

COTS Commercial Off The Shelf

DOD Depth of Discharge

ECI European Components Initiative

ECSS European Cooperation for Space Standardisation

EOL End of Life

EPS Electrical Power Subsystem

ESA European Space Agency

FM Flight Model

FoV Field of View

G/S Ground Station

LEOP Launch and Early Orbit Phase

LEO Low Earth Orbit

MCC Mission Control Centre

NEA Noise Equivalent Angle

NS NanoStar

OBC ON-Board Computer

OBS Organization Breackdown Structure

PCC Payload Control Centre

PDR Preliminary Design Report

PFM ProtoFlight Model

QM Qualification Model

S/C Spacecraft

SDS System Data Summary

SEE Single Effect Event

SM Structural Model

SSO SunSynchronous Orbit

NOMENCLATURE

STR Star Tracker

TCS Thermal Control Subsystem

TID Total Ionising Dose

TMM Thermal Mathematical Model

TRL Technology Readiness Level

UC3M Universidad Carlos III de Madrid

UHF Ultra High Frequency

WBS Work Breakdown Structure

XI

Applicable documents

ID Code Title Issue[AD 01] ECSS-E-ST-10C System engineering general requirements Rev. 1[AD 02] ECSS-E-ST-10-02C Verification Rev. 1

[AD 03] ECSS-E-ST-10-12Methods for the calculation of radiation received and itseffects, and a policy for design margins

Rev. C

[AD 04] ECSS-E-ST-31C Thermal control general requirements Rev. C[AD 05] ECSS-E-ST-60-20C Stars sensors terminology and performance specification Rev. 2[AD 06] ECSS-E-HB-20-07A Electromagnetic compatibility handbook Rev. 1[AD 07] ECSS-Q-ST-70 Materials, mechanical parts and processes Rev. C[AD 07] ECSS-E-ST-10-03C Testing Rev. 2[AD 08] ECSS-E-ST-60-30C Satellite attitude and orbit control system (AOCS) requirements 08/2013[AD 09] NANOST-REQ-070 High level space mission requirements for Phase 1 Rev. 2

Table 2: Applicable Documents.

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[13] David Messmann, Felipe Coelho, Philipp Niermeyer, Martin Langer, He Huang, and Ulrich Walter.Magnetic attitude control for the move-ii mission. In 7th European Conference for Aeronautics andSpace Sciences (EUCASS), Milan, 2017.

[14] E Mooij and DI Gransden. Quasi-transient stability analysis of a conventional aeroelastic launchvehicle. 2017.

[15] P Müller. Esa tracking stations (estrack) facilities manual (efm). Reference: DOPS-ESTR-OPS-MAN-1001-OPS-ONN, 2008.

[16] NASA. State of the Art Small Spacecraft Technology. Small Spacecraft Systems Virtual Institute,California, 2018.

[17] A Pignède. Detumbling of the ntnu test satellite. Project thesis, Norwegian University of Scienceand Technology, Department of Engineering Cybernetics, 2014.

[18] Vincent L Pisacane. The space environment and its effects on space systems. American Institute ofaeronautics and Astronautics, 2008.

REFERENCE DOCUMENTS

[19] Gustavo Emanuel Santos Dinis. Sun-synchronous satellite simulator: An openmodelica simulator.2017.

[20] James A Schoenster and Harold B Pierce. Comparison of vibrations of a combination of solid-rocketlaunch vehicle and payload during a ground firing and launching. 1975.

[21] John Springmann, James Cutler, and Hasan Bahcivan. Magnetic sensor calibration and residualdipole characterization for application to nanosatellites. In AIAA/AAS Astrodynamics SpecialistConference, page 7518, 2010.

[22] James R Wertz, David F Everett, and Jeffery J Puschell. Space mission engineering: the new SMAD.Microcosm Press, 2011.

[23] Cen Zheng, Gao Ruo-fei, and Shen Lei. Lm-3a series launch vechicle user’s manual, 2011.

XIV

1. Introduction

1.1 Objective and Scope

The main objective of this Preliminary Design Report (PDR) is to describe and justify the preliminarydesign solution achieved by Star Worms UC3M team that satisfies customer’s requirements. For thispurpose, the performed concurrent design process is described as well as the engineering models utilised.The IDM-CIC concurrent engineering tool is used as a basis to track the design evolution, including theCommercial Off The Shelf (COTS) equipment list, system budgets and configuration. Hence, its mainoutputs are present in this report.

1.2 Structure of the Document

The PDR is divided in the following sections:• Mission Overview, Requirements Flow-down and Tests:

– Mission Overview: a general description of the mission and the achieved solution is stated.In addition, the state of the art and background of similar missions is included.

– Project Management: definition of the Work Breakdown Structure (WBS).– Requirements Flow-down: the top level requirements are enumerated. Then, the sys-

tem/mission level requirements derived from them and the subsystem level requirementsderived from these previous ones are listed.

– Test plan: delineation of proposed acceptance tests to allow nanosatellite’s customer tovalidate the achieved solution.

• Systems and Concurrent Engineering: description of the iterative process carried out to achievethe optimum solution.• Subsystems Analysis and Design:

– Mission Analysis: detail of the orbit selection trade-off.– System Architecture: description of the architecture of the achieved design with the inter-

connections between S/C subsystems.– Payload, Operational Modes and S/C Subsystems: detail of the subsystem design and

sizing process, engineering models, assumptions and COTS trade-offs.• Risk Analysis and Mitigation: list of detected risks during the design process and the mitigation

actions taken.• Concluding Remarks: the conclusions of the design process are summarised in this section as

well as the proposed future works.• Appendices: System Data Summary, project management charts, Link Budget model charts, TCS

results charts, etc.

1.3 Team Composition and Roles

Alvaro Sanz Casado leads the team as project manager and systems engineer. He is responsible forthe project planning to ensure that the project results are on time and scope with the required level ofquality. Thanks to his work on PROBA-3 ESA mission (Sener Aerospace) and MARIO CubeSat mission(Politecnico di Milano), he has acquired experience in systems engineering. Hence, he has organised thegroup meetings to ease the design iterative process simulating a Concurrent Design Facility and individualmeetings with the rest of team members to provide support in subsystem/orbit geometry design. Finally,Alvaro S. has assessed the environmental effects on the S/C.

Miguel Renieblas Ariño is responsible for the mission analysis and has the role of Communications andCommand and Data Handling (C&DH) Engineer. Miguel has analysed and iterated the different orbitgeometries proposed by the team in the kick-off meeting to achieve the best solution. In addition, he has

Chapter 1. Introduction

developed the communications strategy and the Link Budget between Earth and Satellite. Furthermore,the he has ensured that the generated payload data is stored and transmitted adequately.

Sergio Sarasola Merino is responsible for the Electrical Power System. Sergio controls the PowerBudget and the Power Operational Modes and sets the power generation strategy. Thanks to his studiesin Electronics Engineering, he has been in charge of assessing the electrical compatibility between thedifferent COTS. On the other hand, he has managed the IDM-CIC together with Alvaro Sanz generatingthe System Power Budget, System Mass Budget, S/C CoG, S/C inertia properties and system configuration.

Carlos Álvaro Arroyo Parejo is the ADCS and Structure Engineer. His work is devoted to the computa-tion of the disturbance torques for the given space environment. On top of that, he is responsible for theattitude control and determination strategy taking also into account the S/C attitude requirements alongthe orbit. Besides, C.Alvaro is responsible for the mechanical design and the structure.

Miguel Muñoz Lorente is the thermal control engineer of the mission. His studies in Energy Engineering(Bachelor’s Degree) and Industrial Mathematics (Master’s Degree) have helped him develop a ThermalMathematical Model (TMM) that assesses the evolution of S/C components temperature as a function ofS/C position and attitude in the orbit.

The Star Worms UC3M Organization Breakdown Structure (OBS) is summarised in Figure 1.1.

Figure 1.1: Star Worms UC3M Organization Breackdown Structure.

2

2. Mission Overview, Requirements and Test

2.1 Mission Overview

2.1.1 Mission Statement

The proposed mission is a CubeSat mission to an Earth orbit that shall verify the survivabilityin space of Roscoff worms with the objective of assessing their potential capability of creatingartificial ecosystems for future deep space exploration missions.

2.1.2 Mission goals, Payload and Predesign Ideas

In order to accomplish the scientific objective delineated in the mission statement, the proposed designsolution has to carry a minimal scientific payload that consists of a closed environment artificial ecosystemthat is able to maintain alive, monitor the metabolism and retrieve data of the worms colony.

The mission utilises a CubeSat that carries the biological payload which includes all the systems tomaintain alive the self-sustained colony of Roscoff worms. It deals with the quality and flow of artificial-sea water, the levels of oxygen and carbon dioxide, the quality of the light spectrum and the quantityof photons released for the photosysthesis of the "animalgae", the regulation of temperature and theregulation of the associated microbiome. For this purpose, the payload is equipped with a set of probes, apumping system, an artificial light provided by means of a 1W-LED, a full HD resolution camera andtemperature control devices.

In order to develop a solution which can accomplish this ambitious scientific mission, a concurrent designstudy is performed. Following the CubeSat philosophy, the objective is to find the most robust, maturedand reliable solution that complies with the payload needs and requirements. This implies to minimise thecomplexity of the orbit geometry, system design and relying on matured technologies if possible (i.e useof TRL9 COTS for CubeSat components). The CubeSat philosophy leads to significant cost reduction indetailed design, procurement, assembly, testing, launch and S/C operations. The iterative design processcarried out by the Star Worms UC3M team has led to an optimum solution in terms of orbit selection andsystems design and sizing.

Nevertheless, before starting the iterative process, a brainstorming was performed in terms of rough designconcepts always with the objective of reducing mission/systems complexity. Based on the first missionstatement provided, the following list contains the main predesign ideas that arose in the kick-off meeting:• Cubesat size: since the payload volume is just 3U, can the platform be a 6U structure?• Design choice: There is not a requirement in terms of orbit geometry and launcher availability

has not to be considered. Can the CubeSat avoid equipping a Propulsion System? Because of thedifficulties of miniaturisation, it is the least matured system for CubeSats [16].• Communications: can the Communication system rely only on UHF omnidirectional antennas for

PL data communication?• Power generation: can the EPS system provide enough energy avoiding the use of deployable solar

arrays?• Orbit selection: does a carefully selected Sunsynchronous orbit avoid eclipses while compliying

with the maximum deorbit time? With this, steady power generation capacity and less variantthermal environment would be achieved.

In order to answer/validate those proposed design concepts/solutions both a state of the art study andrough numerical analysis were performed.

Chapter 2. Mission Overview, Requirements and Test

State of the Art

In this section a set of similar already existing mission and mission concepts are described. In this context,the word "similar" means a commonality in expected CubeSat size, PL, orbit, etc.• E. coli AntiMicrobial Satellite (EcAMSat): 6U Cubesat developed by NASA deployed from the

ISS. It carries a 2U biological payload science objective is "to investigate the effect of microgravityon the resistance of the uropathogenic strain of Escherichia coli (UPEC) to an appropriate antibioticand the role of a gene previously identified with respect to antibiotic resistance in this bacterium"[3]. It equips 2 communications modules, one S-Band and one UHF and a dedicated payloadprocessor. It does not have a propulsion system.• TechSat 1: 6U CubeSat developed by Pumpkin, Inc which unfortunatelly was lost in the maiden

flight of Super-Strypi rocket. Its payloads summed a total of 3U volume including a 2U camera. Theobjective of the mission was technologies demonstration. PL data was planned to be transmittedthrough S-Band system while S/C telemetry was spread worldwide continuously with the UHFomnidirectional antenna. In addition, it carries an Inter-Satellite link antenna. It did not have apropulsion system.• GeneSat-1: 3U CubeSat developed by NASA carrying a 2U biological payload whose main objec-

tive was "the study of the effects of the microgravity environment on biological cultures (bacteria,genetic and biological probes to detect “gene expression”)" [9]. Despite the low space available forthe satellite bus (1U), it had an S-Band module, an UHF module and a PL dedicated processor. Itdid not have a propulsion system.

Other missions and concepts were analysed but in general, the common denominator is that those CubeSatsgenerating a significant amount of data rely on S-Band communications and dedicated PL processor forPL data transmission and processing and a UHF module for telemetry. Besides, if the orbit geometry isnot a requirement, existing CubeSats are placed in LEO and do not equip primary propulsion systems. Inaddition, as a rule of thumb, it can be concluded that the S/C components (without the payload) occupyabout 50-66% of the total CubeSat volume. Finally, the answer to the last two posed questions is explainedin the EPS and mission analysis sections respectively.

2.1.3 Summary of the Design Solution

The mission design solution is summarised in Table 2.1. The reader has to take into account that thedescription already includes design decisions taken when designing in detail a S/C subsystem. Thosedecisions are detailed and justified in the sections devoted to that subsystem. Furthermore, the interactionbetween the different subsystems is explained in Section 4.2.

The mission implements a carefully analysed orbit design: The S/C is introduced on a SunsynchronousLow Earth Orbit. With this, firstly, the eclipse periods are minimised which allows a more uniform thermalenvironment along the mission. Secondly, the power generation (steady power generation capacity) andpower distribution strategies are more robust and less complex. Thirdly, the use of a propulsion systemis avoided which generally has the lowest maturity level among CubeSats COTS [16] and implies animportant risk for a biological payload (risk of explosion). And last but not least, end-of life regulation isto be accomplished naturally by interaction of the S/C with the atmosphere.

The solution of the performed concurrent engineering study is a 6U Cubesat (2Ux3U) with 16kg mass(including unit and system margins). Steady power generation is achieved with a set of 3 solar arraysand eclipse-free orbits. An accurate pointing is achieved by means of an star tracker and reactionwheels (3-axis stabilised). The CubeSat is also equipped with a set of fine sun-sensors to allow a quickattitude determination after detumbling and in safe mode. Reaction wheels are desaturated utilising a3-axis magnetorquer. The payload has a dedicated high speed on-board computer to process, compressand storage HD photo/video. Communication with Earth is achieved through a reliable UHF antenna

4

2.1 Mission Overview

and transponder. Payload data is sent to the ground stations by means of an S-Band Antenna and itscorresponding transceiver. A semi-active thermal control strategy with advanced coatings and heaters isconsidered for units’ temperature control. Finally, a lightweight structure and radiation shielding containall the components.

Key ScienceQuestion

- Are Roscoff worms a suitable biological resource to support astronauts’ life in spaceshuttle or space habitat in hostile environments?

Sciene Objec-tive

- Demonstrate the the capability of Roskoff worms to survive in the space environment.

Orbit Geome-try

- Science Operations ∼3 months- Sunsynchronous Low Earth Orbit- Altitude ∼500 km- Inclination ∼97,5 deg- Eccentricity ∼0 deg- Launch Window ∼from FEB to JUL (for no eclipse along the Science Ops.)- Ground Station Coverage per day ∼75 mins

Payload

- Biological Payload- 7 kg-3U-L shaped PL- 9.5 W maximum power consumption- 1 W-LED for worms’ artificial light- Probes, pumping system and temperature control system to simulate and monitoriseworms’ environment- HD resolution camera

Science Prod. - 550 MB data generation per day: HD photos and videos

Spacecraft De-sign

- CubeSat size 6U with <16 kg mass- Communications: UHF deployable antenna and UHF transceiver for telemmetry, trackingand commands; S-Band patch antenna and S-Band transceiver for PL data transmission (2Mbps).- Ground Segment: ESA CORE Network.- OBC: dedicated high-speed-high-storage payload OBC; S/C OBC for satellite houskeep-ing and GNC.- EPS: steady power generation using solar array (∼44 W) and eclipse-free orbits. Batteryto cover power peaks during communication periods.- ADCS: 3-axis stabilised S/C by means of 4 reaction wheels and magnetorquers fordesaturation; star tracker for high pointing accuracy to maximize power generation andfine sun sensors for system redundancy and safe mode.- TCS: semiactive thermal control strategy by means of heaters and advanced coatings.- Structure: lightweight structure with radiation shielding.

Table 2.1: Mission Fact-sheet.

2.1.4 Mission Phases

The successive mission phases and S/C top level operations are explained in this section, starting fromthe launch phase and finishing with the S/C disposal. Before launch, other mission phases exist: missionfeasibility, preliminary design, detail design and S/C manufacture and assembly. However, they are notconsidered in this section that is focused on orbit operations and disposal.

• Launch and Early Orbit Phase: The CubeSat is mount in stack configuration in the launchadapter. During the launch phase, the S/C has to withstand the vibration loads. Then, the launcher

5

Chapter 2. Mission Overview, Requirements and Test

injects the S/C in the target initial orbit (Sunsynchronous orbit, see Section 4.1) introducing a shockload in the S/C. Once separated from the launcher, the next step is to perform the de-spin manoeuvreand deploy the solar arrays. Then, communication with ESA CORE Network is performed andsystems’ status is checked. It lasts about 2 days.• Commissioning Phase: The general health of the CubeSat is checked (main components) as well

as payload status. Besides, small attitude corrections are performed to maximise power generation.Afterwards, the PL control centre commands the payload to initialise the experiments. Finally, thecorrect functioning of all power operational modes is verified. It lasts about 1 week.• Routine Operations Phase: nominal science operations, pursuing the science objectives. S/C

telemetry is transmitted to Earth. PL data is transmitted to the ESA CORE Network through theS-band high data rate communications module. G/S performs S/C tracking and S/C takes this datatogether with the ADCS sensors data to correct the attitude and switch to the appropriate poweroperational mode. The nominal duration of this phase is 3 months but this could be extended inaccordance with the PCC and the MCC .• S/C De-commissioning Phase: the main and most important activity carried out in this phase is

the bacteria/virus elimination command to the payload. The achievement of this target is checkedby the system. Then, the de-commissioning of the rest of subsystems is performed.• S/C Disposal: Once the S/C has finished its operations, it reenters THE Earth’s atmosphere and

burns. This operation is performed by natural decay because of atmospheric drag before the required25 years.

LEOP~2 days

Commissioning Phase~1 week

S/C De-commissioning~3 days

S/C Disposal

<25 years

Routine Operations Phase

3 moths

Mission top level operations break-down

Mount CubeSat in launcher

adapter

Launch phase: withstand

vibrations loads

Launcher separation and orbit

injection: survive to shock

loads

Sunsynchronous initial orbit

De-spin and orientation

Deployment of solar arrays

Establish communication with

ESA CORE Network

System status check

General S/C health is

checked

PL health is checked

Regular communication with

G/S and control activities

Attitude corrections

Participation of the PL

control center to initialise

the PL experiments

Verification of the correct

functioning of the different

power operational modes

Science objectives are

pursued

Steady Power generation

S/C telemetry data

transmission to Earth

PL Data transmission to

Earth

PL command

Attitude corrections

System status check

Potential extension of

Science operations

PL decommissioning:

send command to

eliminate

virus/bacteria of the

biological PL.

Ensure PL

decomissioning

Last system status

check before S/C

decomissioning

S/C end of life

S/C reentry in

Earth’s atmosphere in a

controlled

manner

25 years (max)

safe reentry

trajectory

Figure 2.1: Top Level Operations Breakdown.

2.2 Mission EnvironmentIt is important to understand the environment where the nanosat will operate, as this will have a directimpact on its performance and on its design decisions. The information in this section has been takenfrom "The Space Environment and its Effects on Spacecraft Systems, Pisacane" [18].

As the satellite will be orbiting the Earth, the main concerns are the Earths radiation belts, also known asVan Allen belts. These form due to the Earth’s magnetic field, as solar wind charged particles get trappedin the magnetic field lines, and start moving along them. At the macroscopic level, this generates twobelts with trapped oscillating particles. These regions must be avoided, as they contain a high density of

6

2.3 Project Management

highly energetic charged particles, which can cause damage to the solar cells and the electronics, as wellas to the payload. Therefore, the satellite will have to be placed in Low Earth Orbit (LEO), away from theVan Allen belts.

Here in LEO, there are several factors and phenomena that must be considered:• Earth’s ionosphere: the ionosphere is the ionised part of the Earth’s atmosphere. Its state depends

on many factors, such as the solar activity, and will cause the delay and attenuation of electromag-netic waves used for communications. Therefore, this must be accounted for when designing thecommunications subsystem. Ionised particles in the ionosphere will cause the spacecraft to chargeup. If the spacecraft is not properly designed and built, there will be voltage differences betweensatellite parts, causing high energy discharges that can damage the components. Furthermore, asthe spacecraft will be charged, it can attract sputtered material, creating further surface damage andcontamination.• Atmospheric drag: at lower altitudes, as the density of the atmosphere increases, so will the

atmospheric drag. This will slowly reduce the orbits altitude. This is useful for THIS mission, as itcan be used to de-orbit satellites in a passive way.• Earth’s albedo: the reflection of the Sun’s light on Earth. Second most important external heat

load on LEO spacecraft. This radiation is approximately one third of the radiation incident from theSun. This, together with the infrared heat radiation of the Earth must be considered for the designof the thermal control subsystem.• Space debris and meteoroids: Earth’s LEO is highly populated with space debris. Debris could

impact the spacecraft at extremely large velocities, meaning a small piece of material that cannot betracked from the Earth can render the whole spacecraft unusable.

2.3 Project Management

The project management methodology is a hybrid between the new Agile techniques and classic Require-ments Engineering (V-cycle framework):• Agile Methodology: daily 10 minutes meetings have been scheduled through Hangouts. In those

meetings, every member of the team exposed their achievements of the previous day, the objectivesfor the present day and the obstacles for continuing their work. With this, the wasted effort andre-do times have been minimised. On top of that, a larger meeting was programmed every week tomonitor the overall progress of the project.• Requirement Engineering: the lower level requirements have been derived from the top level require-

ments provided in the mission statement following the ECSS-E-ST-10C in a V-cycle framework.

In addition to the technical meetings stated in Section 1.3, four key meetings were held:• Kick-off meeting: project initiation. Mission statement reading and understanding. Literature

review and first predesign ideas and concepts. Establishment of the project scope, WBS (FigureD.1), Work Flowchart (Figure D.2) and project time-line. Understanding of the CubeSat conceptand design philosophy.• Requirements Review meeting: review of the derived lower level requirements from the top level

requirements.• Conceptual Design Review: review of the physical soundness and coherence of the design concept.

Correction of potential mistakes. Check requirements compliance.• Detailed Design Review: review of the achieved solution. Take action to correct mistakes. Check

requirements compliance.

From the technical point of view, concurrent engineering has been performed to iterate the design (seeChapter 3). To ease this process, the IDM-CIC software has been used. Besides, the System DataSummary was created (see Appendix H). The SDS is a typical industry document that includes all therelevant information of a determined component: mass, TRL, power consumption, input voltage, overall

7

Chapter 2. Mission Overview, Requirements and Test

dimensions, thermal properties, electric interface, performance parameters, etc. This document (GoogleSheet shared in Google Drive) allowed all team members to have a clear and easy access to selected COTSinformation enhancing the iterative concurrent design process.

2.4 Requirements FlowdownThe mission requirements are listed in Table A.1. The majority of them have been derived directly (seeTable A.2) from the information contained in the mission statement provided by the customer. Otherrequirements have been derived from the mission analysis and design iterations (i.e NS-SYS-12). Thethird column of the table indicates the parent requirement for each listed requirement. Besides, the fourthcolumn indicates the verification method to such requirement based on ECSS-E-ST-10-02C. For thoserequirements verified by analysis, the fifth column indicates in which section they are addressed. Thereare three levels of requirements:

1. Top Level Requirements: there are 10 top level requirements listed in the mission statement.2. System/Orbit Level Requirements: they are derived from the Top Level Requirements. They

consider aspects involving the whole system or the mission analysis3. Subsystem Level Requirements: they are derived from the System/Orbit Level Requirements and

affect a specific subsystem.

The derivation of lower level requirements from the top level requirements ease the verification ofthem. However, in order to provide some guidelines of how and/or where the top level requirements areaddressed, a summary table is included in this section (Table 2.2).

2.5 Acceptance TestsIn the Phase D of the programme, a test campaign shall be performed. By that time, the payload willhave passed its Critical Design Review, so it will be much more characterised. The payload (as the restof CubeSat components) will be qualified (through unit level testing) for sine vibration levels, randomvibration levels, shock levels, etc. Hence, when integrating the PL on the structure, the platform responsibleshall ensure that those levels (with margin) are not surpassed. According to ECSS-E-ST-10-03C, testswith different components models (SM, QM, PFM and FM). Table 2.3 lists some of the first tests basedon the ECSS and experience that shall be performed to allow nanosatellite’s customer to validate theachieved solution. In these tests, a PL model is integrated in the platform to verify certain aspects.

8

2.5 Acceptance Tests

ID GuidelinesRW1 Several requirements have been derived from RW1 in order to cover all the different

aspects involved in it: power, thermal control, configuration, C&DH and mission analysisrequirements (covered in the corresponding sections).

RW2 This requirement is comprised in the coverage analysis and the comms. link budget.RW3 This requirement is covered in the TCS section and the S/C shielding analysis. However,

RW3 encompasses more aspects that shall be covered in more detailed design phases (PhaseC): TID analysis, S/C outgassing, S/C contamination, etc.

RW4 This requirement is covered in the ADCS section.RW5 The final configuration is a 6U CubeSat.RW6 This requirement is covered in the mission analysis section, but affects other aspects as the

EPS (solar array degradation) or the TCS (non-normal thermal cases).RW7 This requirement is covered in the mission analysis section.RW8 Only ESA CORE Network is considered in the S/C Coverage analysis.RW9 This requirement is covered in the C&DH and Structure sections. Besides, during the

project lifecycle, the PL will be manufactured, manipulated, integrated and tested followingthe required procedures to avoid any biological contamination.

RW10

This requirement is covered by the review of design:- ROHS and REACH: aspects related to the use of chemicals or other hazardous substances.The detail design (Phase C), manufacturing, assembly and testing processes (Phase D)will comply with this regulations.- CSR: a stakeholder analysis will be performed in the mission definition phase: to whomand how does this mission affect (directly or indirectly). Besides, all the ProgrammePhases will be carried out minimising their environmental impact and being socially andeconomically responsible. In order to accomplish these last objectives, specific analysisand evaluations will be carried out and the actions to be taken will be detailed in a particularControl Plan.

Table 2.2: Top Level Requirements verification guidelines.

ID Name PL Mod. Description/Objective Test Procedure

1Shock 

LevelsSM

Collection of shock levels at PL's mounting interface 

due to every possible cause: stage separation, pyro 

device, etc,

Apply in the Satellite‐CubeSat deployer interface the 

expected shock input and measure the levels at the PL 

interface

2Random 

VibrationSM

Collection of random vibration levels at PL's 

mounting interface ocurred in the launcher: 

aerodynamic excitations, ground handling, etc.

Apply in the Satellite‐CubeSat deployer interface the 

expected random vibration levels over a freq range of 20Hz to 

2kHz and measure the levels at the PL interf.

3Sine 

VibrationSM

Collection of sinusoidal vibration levels at PL's 

mounting interface ocurred in the launcher: 

coupling of launcher frequencies, imbalances in rot. 

elements

Apply in the Satellite‐CubeSat deployer interface the 

expected sine vibration levels and measure the levels at the 

PL interface

4

Full 

Functional 

Test

QM

Test that demonstrates the integrity of all functions 

of the item under test, in all operational modes and 

all foreseen transitions

Command the S/C: power ON/OFF, switch among op. modes, 

activate PL, end PL experiment, activate comms module, test 

failure modes, etc.

5Environm. 

TestQM

Tests applied to a product simulating environmental 

conditions as encountered during its op. life cycle

Apply the expected thermal environment to the S/C to 

correlate the analytic models

6Polarity 

TestQM

Verify the correct polariry of the functional chains 

(mainly PL and AOCS)

Command the units and verify that the response happens as 

expected. i.e Payload water rotation in the desired direction

7Acoustic 

SpectrumSM

Max. value of the time average r.m.s. sound 

pressure level in each freq. band occurring inside 

the CubeSat deployer

Subject the CubeSat to the maximum expected acoustic 

spectrum, specified in octaves over a frequency range of 

31,5Hz to 10kHz and measure the response

Table 2.3: Acceptance Tests Proposal.

9

3. Concurrent Engineering

In this section, a brief explanation of the Iterative Concurrent Design is given. The reader has to take intoaccount that only the key points are explained in this chapter and further details are given in chapter 4.

Before starting the Concurrent Design process itself, the scope of this preliminary analysis was clarifiedwith the customer in the first Slack meeting based on the first issue of the High Level Requirementsdocument of the mission. Then, in the team’s kick-off meeting some rough mission concepts wereproposed and the state of the art was studied (see Section 2.1.2). Then, to ease the analysis of how a S/Cfunction affects each other, a N2 chart (see Figure B.2). This chart includes the different S/C subsystemsand mission segments and sets their functional relationship. Afterwards, the Concurrent Design processstarted in the framework of the top level requirements included in the mission statement and the derivedrequirements detailed in Section 2.4. Hence, summarising, the starting point is a CubeSat that: shall nothave propulsion system, shall be introduced in a SSO and shall be capable of transmitting the generatedPL data per day in one day.

Figure 3.1 depicts the iterative Concurrent Design flow. PL data (mass, shape and power) are given in themission statement. Besides, a first estimation of the TCS power consumption based on rough analysis,experience and state of the art was set. These values are fixed for the subsequent iterative cycles. Forthe first iteration, an estimation of mass, inertia, configuration and shape was performed on top of thefixed values. This is the reason why this boxes are marked with a "1". The same applies to the first SSOaltitude guess.

Firstly, with the altitude guess, an eclipse analysis is performed to check whether its duration requirementis satisfied. If the answer is positive, a coverage time with the ESA CORE Network for the defined orbit iscomputed. Besides, the de-orbit time as a function of mass is estimated. On the one hand, the coverageanalysis together with the PL data transmission requirements sets the minimum required Data Rate. Thisdetermines the Communications and C&DH subsystems sizing (including the link budget). On the otherhand, the orbit (external torques for the guess orbit) in conjunction with the ADCS requirements and themass, inertia and S/C configuration condition the ADCS sizing.

After the subsystem sizing, a COTS market research is carried out, since the objective is to rely only onCOTS components. Several trade-offs, which are detailed in their respective Section 4, are performed.There are two main outputs of the trade-offs:

1. ADCS, Communications and C&DH components power consumption, which sizes the EPStaking also into account the subsystem power consumption fixed values (PL and TCS). Once again,a COTS market research (in this case for the EPS) is performed. The outputs of this branch are theS/C power budget and S/C power generation strategy (based on the power operational modes).

2. ADCS, Communications and C&DH components mass and size, that in combination with theoutput of the EPS trade-off and the fixed PL properties update the S/C mass budget. Then, S/Cconfiguration is built and CoG and inertia properties derived.

With the new mass estimation and the S/C configuration, the de-orbit time for the given SSO is computed.It is important to remark that both power and mass calculations include the required margins dictated bythe ECSS. Finally, the two questions that close the loop are answered:

1. Does the design solution achieved fulfil the requirements?2. Has the design iterative process converged?

If the answer to both questions is affirmative, the design is valid and the iterative loop can be stopped. On

the contrary, a negative answers implies beginning another loop. Nevertheless, from the second loop, theS/C Power Budget, S/C Mass and Inertia and orbit altitude guesses are updated based on the results of theprevious loop rather than using the first rough estimations of the first iteration. With this, convergence isenhanced.

In any case, a valid design solution can still be optimised. After reaching a valid design solution, StarWorms UC3M team performed a design optimisation in terms of S/C power consumption and data rate(through coverage time).

Design Estimation

TCS power

Design Choice

550 MB PL data per day transmmited

every day

PL power

PL Mass & shape

Initial SSO study f(h)

ADCS sizing

Coverage time De-orbit time=f(m)

Altitude guessEclipses

period OK?

No

Yes

External torques=f(h)

Data Rate

S/C PowerBudget guess

SA mass and shape guess

Mass and Inertia guess

1

1

COMS and C&DH sizing

COTS Market Research

COTS Market Research

COMS and C&DH power

COMS and C&DH mass and shape

ADCS mass and shape

ADCS power

EPS sizing

COTS Market Research

S/C Mass, Inertia and

Configuration

S/C power BudgetS/C power gen. strategy S/C De-orbit time

1. Does the design solution achieved fulfil the requirements?2. Has the design iterative process converged?No

1

Yes

Valid Design

This symbol means that the value inthe 1st iteration is estimated frompartial data and/or best engineeringjudgement

1

Figure 3.1: Concurrent Design Flowchart.

11

4. Subsystems: Analysis and Design

4.1 Mission AnalysisThe spacecraft must be in an orbit where it can guarantee the survivability of the payload, that does notrequire any kind of propulsion to accomplish the requirements and where it can have enough coveragetime to transmit all the data produced every day.

4.1.1 De-Orbit AnalysisThe most limiting requirement for orbit selection, given that there is no propulsion system, is requirementRW7, as then the satellite must be able to de-orbit in a passive way through the use of atmospheric drag.This can be modelled by applying a force opposite to the velocity of the satellite. This force is proportionalto the square of the velocity of the satellite, its wet surface area relative to the flow, and to the air density.As air density drecreases with altitude, the higher the altitude, the lower the drag force is, meaning athigher altitudes, the spacecraft will take more time to de-orbit. Therefore, in order to be able to de-orbitthe spacecraft in 25 years, there is a limit in the altitude of the orbit at the start of the mission. This limitis determined by the ballistic coefficient of the spacecraft (the ratio of mass to surface area). Therefore,the worst case scenario is a spacecraft with the smallest possible surface area, and the highest possiblemass. A de-orbit simulation was performed using STELA for this worst case scenario.

0 100 200 300 400 500 600

0

5

10

15

20

25

30

35

Figure 4.1: Re-entry simulation for a surface area of 0.02 m2.

As seen in Figure 4.1 from the simulation, it can be determined that the maximum altitude of the orbit inorder to comply with requirement NS-ORB-03 is around 520km. An increase in mass causes an increasein de-orbit time. So a conservative approach must be taken for the estimation of the mass to allow forerrors and ensure the requirements are met.

4.1.2 Eclipse AnalysisIn order to maximise power production for payload housekeeping and for communication, eclipse timeper orbit must be minimised. The solution to this is to use a dawn-dusk sun-synchronous orbit (SSO).This kind of orbit takes advantage of the J2 effect, which has to do with the fact that the earth does nothave a spherical mass distribution. This causes the right ascension of the ascending node (RAAN) to driftwith time depending on the height and inclination of the orbit. For a given orbit altitude, it is possible

4.1 Mission Analysis

to select an inclination such that the change in the rate of change of the RAAN is equal to the angularspeed of the Earth around the Sun. This orbit can be seen in Figure 4.2. Here the orbit plane, shown inred, rotates at an angular rate ω equal to that of the Earth around the Sun.

Figure 4.2: Representation of a sun synchronous orbit.

This orbit was then simulated over a whole year at different altitudes using GMAT, and the results areshown in Figure 4.3. In the figure, it can be seen that there are only Earth eclipses during one part ofthe year, and that increasing the altitude decreases the eclipse time and the period where eclipses occur.Additionally, it can be seen that if the mission is launched between the months of February and July, thespacecraft will not experience any Earth eclipses during the operating part of the mission. Increasingthe altitude would increase this ‘no eclipse’ launch period. Having no eclipses implies a higher powergeneration and allows for a simpler and more efficient thermal control subsystem, as hot-cold cycles arereduced. There are two Lunar eclipses in summer at the same day of 20 minutes each, therefore this shallbe taken into account for the thermal subsystem design as the worst ’cold case’ scenario.

0 50 100 150 200 250 300 350

0

1

2

3

4

5

6

7

Figure 4.3: Eclipse time in an orbit along the year for different altitudes.

13

Chapter 4. Subsystems: Analysis and Design

4.1.3 G/S Coverage Analysis

Another thing that must be taken into account to decide on the final orbit is the coverage time. This mustbe as large as reasonably possible as it sets the data rate requirement. Coverage analysis for the ESAground station CORE network, shown in Figure 4.4, was performed using GMAT and STK simulations,where it was assumed that the minimum elevation at which the satellite is seen by any ground station is 15degrees. The results obtained by both programs do not take into account that the satellite can be seen bytwo or more ground stations at the same time, giving the sum of coverage time as if they were isolatedcontacts (see ground station contacts in Figure E.1). It is important to consider this, as all of the groundstations in Europe have an overlapping coverage region. Therefore, a specific software was created todetermine the actual contact time taking into account overlapping contacts. A sun-synchronous dawndusk orbit was simulated for 10 days at three different altitudes, and the mean contact time per day wasobtained, this is shown in Table 4.1.

Figure 4.4: Orbit ground track and ESA CORE Network coverage simulation.

Altitude (km) Coverage time (minutes per day)400 34500 76600 130

Table 4.1: Contact time per day for different orbit altitudes.

As seen in Table 4.1, coverage time increases with altitude. These numbers do not take into accountground station availability, so a safety margin must be taken. Therefore, as coverage time and launchwindow increase with orbit altitude, the highest possible orbit is wanted that will still cause the satellite tode-orbit passively in less than 25 years.

The selected orbit after performing the mission analysis was a circular dawn-dusk SSO at an altitude of500km and an inclination of 97.4 degrees to the Earths equator. This is because at 500km, the de-orbitrequirement is satisfied with a safety margin; the launch window is maximised for not having eclipses inthe 3 months and a week of operation; and given the coverage time and the amount of data generated everyday that have to be transmitted given requirement NS-ORB-01 it yields a required data rate achievable bythe communications COTS, as seen in Table 4.2.

14

4.2 Mission Architecture

4.2 Mission ArchitectureThe mission architecture is depicted in Figure 4.5. The reader has to take into account that the descriptionalready includes design decisions taken when designing in detail a S/C subsystem. Those decisions aredetailed and justified in the sections devoted to that subsystem.

The mission architecture could be divided into two segments: the Ground Segment and the S/C Segment.On the one hand, S/C telemetry is received and tracking and control of the S/C (via commands) areperformed in the mission control centre. On the other hand, the scientists of the Payload Control Centrereceive the payload science data and forward PL commands (change photo/video compression rate,modify the number of photo/video per day, etc.) to the Mission Control Centre to send to the CubeSat.Communications in Ground Segment rely on the ESA CORE Network.

Regarding the S/C Segment, there are three subsystems connected between them and to all the remainingones: the EPS, the TCS and the S/C OBC. Firstly, the EPS generates energy through solar arrays to feedall S/C components and charge the battery. Commanded by the OBC, it provides the required power toeach unit as a function of the power operational mode set (safe mode, science mode, communicationsmode, etc.). Secondly, the TCS is in charge of regulating and controlling the temperatures of the differentS/C components either actively of passively.

However, the core of the S/C is the OBC, where all telemetry coming from the rest of subsystems isprocessed and S/C housekeeping is performed. This telemetry that includes information related to unit’stemperature, voltages or currents is forwarded to the G/S through the communications subsystems (viaUHF TX/RX). If the OBC detects an overload, an anomaly or a failure, it changes the operational modeto "safe mode" to protect the S/C. Besides, if a given critical unit is becoming too cold (reaching itsminimum design temperature threshold) the OBC acts over the TCS to switch on the heater. Finally, theOBC switches the above-mentioned power operational modes as a function of the mission phase or S/Cposition in the orbit (tracking received from G/S and propagated in the OBC).

The sensors of the ADCS provide S/C attitude to the OBC. The computer processes this information andelaborates the required control laws to orient the CubeSat to the desired direction. These control laws areforwarded to ADCS actuators which control the attitude. Nominally, the OBC commands the ADCS tokeep solar array sun pointing and antenna facing the Earth.

In the Science Phase of the mission, the payload acquires photos/videos of the Roscoff Worms which aresend to the dedicated payload processor where are processed, compressed and stored. This information isthen forwarded to the Earth by the S-Band module (high data rate) of the Communications Subsystem.

Finally, the structure & bus subsystem serves as platform for the S/C subsystems. In addition, it protectsthem against undesired radiation, and acts as a passive thermal control device, avoiding/enhancing internalheat going out or external heat entering the S/C.

15

Chapter 4. Subsystems: Analysis and Design

Mis

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TCS Power + Excess

Power→

←EP

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Subsysttemp →

← OBC Temp ctrl

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PL ProcTemp ctrl

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Figure 4.5: System Architecture.

4.3 Payload

The payload in this mission is the main scientific load, devoted to study the behaviour of a Roscoff wormcolony in outer space conditions. For this reason, it is the differentiating element in the development,design and analysis of the mission. Inside its architecture, it contains every unit devoted to the survival ofthe colony (pumps with water and nutrients, light, thermal control, etc.) and the scientific supervision(cameras, sensors...). The PL is considered in this preliminary analysis as a black box with strongconstraints to guarantee the mission objectives. The PL main characteristics are:

16

4.4 System Operation Modes

• Weight and density: 7kg (constant density assumed)• Volume: 3U• Shape: L-shaped• Thermal model: isothermal• Temperature range: 11-15oC• Heating power installed: 2W• Camera power demand: 5 W• Pumping, probes and LED power demand: 2.5 W

4.4 System Operation ModesThe S/C will change among different operational modes along the different mission phases (see Section2.1.4). One of the key points is the power consumption. Table 4.7 shows the power sets of S/C units as afunction of the power operational mode The S/C operational modes are:• Nominal mode: the PL is operating nominally, performing all the science tasks. The rest of systems

are also working nominally, with the S/C pointing towards the Sun with its solar array and theS/C antennas facing the Earth. This is performed with a rolling motion around S/C +X axis (seeFigure 4.10) in which the period of rotation coincides with the orbital period. The dedicated PLprocessor saves PL data on its storage device. The TCS is active and precise attitude determinationis accomplished with the Star Tracker. The unique system that is at idle set is the S-Band module ofthe communications system, waiting for a communication window.• Nominal+Communications mode: all systems are functioning in nominal mode but in this case

the communications module is transmitting PL and S/C data to the G/S.• Safe mode: used in the deployment phase and in a non-nominal situation (internal failure or

extraordinary space environmental event like a coronal mass ejection (CME)). In this case, the S/Creduces the power consumption to the minimum, switching all non-critical S/C systems. OBC,TCS and PL sensors, pumping and heating are ON. In the deployment phase (just after launcherseparation) or if in Sun-pointing attitude is lost, the Sun sensors provide quick and reliable attitudedetermination to rapidly correct S/C’s attitude and recover power generation. In such case, thebattery provides enough energy to feed all S/C systems.

4.5 Communications Subystem (CS) and Ground Segment (GS)The communications subsystem is responsible for the communication between the spacecraft and theground stations. This involves sending payload and telemetry information, as well as receiving commandsfrom the ground station.

4.5.1 Frequency selectionDue to the current nanosat market, there are limitations in the frequencies that can be used for communi-cations, as usually, there are only UHF/VHF and S-band products available. As the data rate is reasonablyhigh, UHF or VHF are not valid to transmit the payload data, so S-band must be used. However, in orderto increase the robustness of the system and ensure that the satellite can communicate at all times witha ground station when in sight, a UHF/VHF is used to send telemetry information and receive groundstation commands. As these communications require much less power, an omnidirectional Antenna canbe used. This configuration has been used extensively for nanosat designs, as seen earlier in Section 2.1.2.

4.5.2 COTS trade-offIn this section, a COTS market research is performed. Several options for each component are presentedand a trade-off according to different aspects with their respective importance/weight (indicated in eachtable) is carried out. The colour of the cell indicates the score that each component has in each aspect: red

17

Chapter 4. Subsystems: Analysis and Design

(1 point), yellow (2 points) and green (3 points). Hence, the colour value is multiplied times the weight ofthe considered aspect and finally all row values are added. The max possible score is indicated in the lastcolumn of the table.

A transceiver and an antenna were selected for each frequency. This was done by performing a trade-offstudy for each of the components, comparing several manufacturers. For transceivers, the parameterswith which the different antennas were compared were: data rate, the power required for transmission,the mass, and the level of detail of information that was given by the manufacturer in the datasheet. Thetrade-off studies for the S-band and the UHF transceivers are shown in Tables 4.2 and 4.3 respectively.

COTS Date rate (kb/s) Tx power (W) Mass (g) Available info TotalWeight 1 2 1 1 Max = 15

GOMspace TR-600 2000 4.13 65.3 Complete 13Endurosat S-Band TX 5000 7.2 250 Limited 10ISIS space TXS 4300 10 120 Limited 8

Table 4.2: Trade-off study for S-Band transceiver.

COTS Date rate (kb/s) Tx power(W) Mass (g) Available info TotalWeight 1 2 1 1 Max = 15

GOMspace AX-100 38.4 2.64 65.3 Complete 13Endurosat UHF TRX 19.2 2.5 94 Limited 11ISIS space UHF TRX 9.6 4 75 Limited 7

Table 4.3: Trade-off study for UHF transceiver.

For the S-band antenna, the parameters are gain, power consumption in transmitter mode, half powerbeamwidth (HPBW), mass, and datasheet completeness. For the UHF antenna, the parameter of HPBWis changed for ease of integration with an S-band antenna, as the UHF antennas commercialised are allomnidirectional, and both antennas will be located at the same face of the satellite. These trade-offs areshown in Tables 4.4 and 4.5 respectively.

COTS Gain (dB) Power (W) HPBW (deg) Mass (g) Available info TotalWeight 2 2 1 1 2 Max = 24

GOMspace 8.4 10.7 66 110 Complete 21Endurosat S-Band ant

7.9 Unknown 71 64 Incomplete 13

ISIS space 6.5 Unknown 100 Unknown Incomplete 10

Table 4.4: Trade-off study for S-Band patch antenna.

COTS Gain (dB) Power (W) Integration Mass (g) Available Info TotalWeight 2 2 1 1 2 Max=15

GOMspace 1.6 0.17 Supported 30 Complete 24Endurosat >0 Unknown Difficult 85 Incomplete 11ISIS space Unknown Unknown Possible <100 Incomplete 9

Table 4.5: Trade-off study for UHF antenna.

The outcome of the trade-offs was that the transceivers and antennas selected are the ones from GOMspace.One key aspect is that the manufacturer provides all the information that affects the overall concurrent

18

4.5 Communications Subystem (CS) and Ground Segment (GS)

design and to perform a RF link budget. This is important as then the calculations at the PDR stage willbe more accurate.

4.5.3 Link budgetTo calculate the link budget, the transmitting and the receiving antennas parameters, as well as environ-mental parameters must be known. The antenna parameters are obtained through datasheets, while theenviromental parameters are obtained through tables and graphs. These parameters will now be defined inorder to have a better understanding of how a RF link budget is produced, and its meaning.

The Equivalent Isotropic Radiated Power (EIRP) is the equivalent power required by an ideal isotropicantenna to produce the same power density as the one observed in the direction of maximum gain of theantenna. It represents the performance of the transmitting equipment. The EIRP is defined in Equation4.1, where GT M,SAT is the gain of the transceiver in dBm, GANT,SAT is the maximum gain of the antenna,and LT M−ANT,SAT are the transmission losses/line losses from the transmitter to the antenna.

EIRP = GT M,SAT −LT M−ANT,SAT +GANT,SAT (4.1)

Path loss (Lpathloss) is the loss due to the propagation of the signal to the receiver, and is composed of freespace losses, rain attenuation losses, and attenuation through atmospheric gases.

Lpathloss = L f ree +Lrain +LAtmosphere (4.2)

Free space loss is the loss between two isotropic radiators in space, which can be calculated with Equation4.3, where D is the distance between the antennas, and λ is the wavelength of the signal. For this case, Dis taken as the distance from a ground station at sea level to the satellite at 15 degrees of elevation withrespect to the ground station, as this will be the maximum distance the satellite will communicate over.

L f ree = 20× log10

(4πD

λ

)(4.3)

Rain attenuation and atmospheric gases losses depend on the frequency of the signal and the elevationangle, and they arise from EM wave absorbtion by water and gas molecules in the atmosphere respectively.These parameters can be estimated by looking at attenuation profiles, such as Figures E.3 and E.2.

The antenna pointing loss is the loss due to the transmitting and receiving antennas not being alignedalong their maximum gain direction. This is estimated through Equation 4.4, from [12], where θ is thehalf power beamwidth of the antenna, and e is the pointing error, or the divergence from the direction ofmaximum gain.

Lθx = 12( e

θ

)2(4.4)

The carrier to noise spectral density ratio represents the link performance, and it is defined in Equation4.5. Where G/T is the gain to noise temperature ratio, which represents the ratio between the gain thatthe receiver could provide and the amount of noise; k is the boltzman constant; and LTotal is the sum of allthe losses, composed of path and pointing losses.

C/N0 = EIRP+G/T − k−LTotal (4.5)

The energy per bit noise ratio (Eb/N0) is the ratio of the received energy per bit, to the noise spectralpower density, and is given by Equation 4.6, where R is the data rate in bits per second. For a desired biterror rate (BER), the probability of receiving one bit incorrectly; there is be a minimum Eb/N0, whichdepends on the error correction code used (QPSK R=1/2 Reed Solomon). This is illustrated in Figure E.4.And then the recovery margin is the difference between the achieved Eb/N0 minus the demodulation lossand the required one. This should be at least 2dB (req. NS-COM-03), in order to account for errors.

Eb/N0 =C/N0−R (4.6)

19

Chapter 4. Subsystems: Analysis and Design

Table 4.6 shows the RF link budget for the S-band downlink. This is the most restrictive case, as for theUHF case, the data rate required is far lower than the one offered by the COTS components. For theS-band downlink, the communications subsystem must be capable to transmitting 2000 Kbps at a distanceof approximately 1400 km with a BER smaller than one in a million. From the table, it can be shown thatthe system is capable of fulfilling the requirements (NS-COM-03, NS-COM-04 in Table A.1), and has areally high margin. The Gain to Noise Temperature Ratio, as well as the pointing error for the groundstation were taken from the ESA CORE Network station datasheets [15].

OrbitEarth Radius 6378.14 kmSemi-Major Axis 6878.14 kmInclination 97.4 degreesElevation Angle 15.0 degSlant Range 1,407.52 km

Spacecraft AntennaTransmission output power 1.58 WTransmission output power 2.00 dBWTransmission Gain 8.4 dBiTransmission Line Losses 1.0 dBTransmission Mod Losses 0.0 dBEquivalent Isotropic Radiated Power 9.40 dBHalf Power Beam Width 66.0 degSpacecraft Depointing 75.0 degfrequency 2250.0 MHzWavelength 0.13 m

PathSpacecraft pointing loss 15.50 dBPolarization loss 1.0 dBPath loss 162.45 dBAtmospheric loss 0.3 dBIonospheric loss 0.1 dBRain loss 0.0 dBIsotropic signal Level at Gound station -169.90 dB

Ground Station AntennaGound station depointing 0.1 degHalf Power Beam Width 0.6 degGound station pointing loss 0.21 dBGain to Noise Temperature Ratio 29.1 dB/KCarrier to Noise Spectral Density Ratio 87.59 dBHzEnergy per Bit to Noise Ratio 24.60 dBHz

Data RateData Rate 2000 KbpsCorrection Code QPSK, R=1/2, Reed SolomonBit Error Rate 1.00E-06 ProbabilityRequired Energy per Bit to Noise Ratio 2.5 dBDemodulator Loss 1.0 dBRequired Margin 2.0 dBAchieved Margin 21.1 dB

Table 4.6: S-band downlink RF link budget.

20

4.6 Command & Data Handling

4.6 Command & Data Handling

The command and data handling subsystem is responsible for receiving and distributing commands toother systems and for collecting and processing payload and housekeeping data. For this mission, aconfiguration with two on-board computers was selected. One of them (the PL OBC) is responsible forthe storage and processing of all the data generated by the payload’s camera, as well as transmitting itthrough the S-band transceiver. On the other hand, the S/C OBC is responsible for:

• Satellite and payload housekeeping and telemetry.• Processing of tracking information to correct/update the orbit propagator inside the computer’s

software.• Execution of commands received from the ground through the UHF transceiver.

A list of the ground station commands that the system may receive is shown below:• Activate the dedicated payload computer and the camera at the start of the mission after the

deployment of the solar panels and the UHF antenna.• Modify ADCS control laws.• Modify camera settings (ISO, camera aperture, video frame rate).• Modify payload settings (photograph and video frequency, amount of compression/quality of file

sent).• Send again a certain picture or video at the desired quality or compression.• Prepare for decommission: terminate payload life according to requirement NS-CDH-03 and

discharge the battery

As the transceivers used are from GOMspace, and this manufacturer also has on-board computers whichsatisfy the requirements, both on-board computers where chosen from GOMspace. By doing so, the riskof not being able to integrate the components is mitigated which will make the integration phase mucheasier, reducing costs. Furthermore, as the components come from the same manufacturer, they can beintegrated into the same motherboards, reducing volume and weight.

The payload computer selected is the GOMspace NanoMind Z7000, which comes with Linux operatingsystem, making it relatively simple to program, and capable of carrying image and video file processingsoftware. It also has a storage capacity of up to 32 GB, meaning it meets the payload data storagerequirement NS-CDH-01. The housekeeping and telemetry computer chosen is the NanoMind A3200.Both of the computers are compatible with the two transceivers.

4.7 Electrical Power Subsystem (EPS)

The EPS is responsible for acquiring, regulating and distributing the power required to all S/C components.As the majority of Earth missions do, the CubeSat is equipped with solar panels, a battery and a PowerManagement And Distribution (PMAD) unit.

The energy obtained by the solar arrays is received by the PMAD. The PMAD is composed by theArray Conditioning Unit (ACU), which manages that received energy and derives it to the other PMADcomponent, the Power Distribution Unit (PDU). The PDU is responsible for distributing the power at thevoltage required by the unit. This PMAD is also responsible for making sure that the maximum currentload that the systems/units can withstand is not surpassed. Finally, a battery is always required to at leastprovide power to the critical units when power generation capabilities are limited.

With the EPS architecture defined, an iterative process to size and select the different component isperformed. As stated in the Mission Overview (Section 2.1) the components shall be COTS. This activityis immersed in the whole S/C concurrent design as stated in chapter 3 (see Figure 3.1). On the one hand,

21

Chapter 4. Subsystems: Analysis and Design

the main inputs to this process are the selected orbit (already set to a 500 km SSO), an estimation of thepower consumption of all S/C subsystems and PL and the operational modes (see Section 4.4). On theother hand, the outputs are the power generation strategy, the solar array size and configuration and thebattery size.

4.7.1 Power Generation StrategyThe SSO orbit was selected to have a steady power generation capability without Earth eclipses (seeSection 4.1). This orbit together with the operating modes allows to have one side of the CubeSatalways oriented towards the Sun (+X direction, see Figure 4.10) minimising (or even eliminating) thecosine losses. Besides, the coverage time that determines the maximum time in nominal+communicationoperative mode was computed in the mission analysis. Taking one day as reference, this time representsonly a small fraction (about 75 mins).

Once estimated the power consumption of the rest of S/C subsystems, the communication subsystemCOTS research (Section 4.5) showed that the required power for the S-Band module was about 1/3 ofthe total estimated power consumption in nominal+communication mode. Since that S-band moduleis not switched ON in nominal operation (22.75 h out of 24 h), it would be a waste of energy sizingthe solar arrays to that peak power of communications mode (see Table 4.7 for the last iterated powerbudget). Furthermore, despite not being a restriction in solar array size, that philosophy would implya significantly larger solar array (at least 12 cells more). Hence, to avoid the energy waste and reducethe solar array size, the optimum solution is using the battery to cover that extra power consumptionduring nominal+communications mode while the solar array is sized to cope with the nominal mode.Taking into account component and system power margins the power to be provided by the battery innominal+communications mode is:

∆Power = Powercom−Powernom = 54.13−38.152 = 14.43W (4.7)

4.7.2 Power Budget and Power Operational ModesTable 4.7 shows the peak, idle and nominal power consumption of S/C components. This table showsthe results of the last iteration of the concurrent design process and values are taken from componentsdatasheets (see Table 4.25 for component list). PL power consumption was taken from mission statement.TCS power consumption was estimated to be 10 W (6 W available in the finally chosen battery and anextra 4 W). TCS power is assumed constant for all modes. The last three columns of Table 4.7 show thepower consumption at each operational mode. It takes the corresponding power value of each component.

This tables do not include any power margin except in the last row (includes component and systemmargin). However, all calculations are performed with component and system margin. Despite themajority of components being COTS, it was decided to apply 10% power margin (conservative) to all ofthem since their datasheets do not always provide the different power consumption (peak, nominal andidle). On top of that, a 10% system margin is added. See IDM-CIC output for more detailed information.

4.7.3 Solar Array and Battery SizingAs it was stated above, the solar array is sized considering the nominal operation mode. However,this design has to account for the power used to charge the battery that will be used later on the nomi-nal+communications mode. To solve this problem, an energy balance is performed considering a periodof 24h. It is important to remark that these calculations are performed using power values with margin.

Egenerated = Econsumed

Egenerated = TlightPEOL(4.8)

Where

22

4.7 Electrical Power Subsystem (EPS)

Power [W] UnitNum.

UnitPeak

UnitIdle

UnitNom

SAFEMODE

NOMMODE

NOM+COMMODE

AOCS 3.26 4.47 4.47RW 4 1.50 0.01 0.60 0.60 0.60 0.60Magnetorquer 1 1.20 0.18 1.20 0.00 1.20 1.20Sun Sensor 3 0.07 0.07 0.07 0.07 0.07 0.07Star Tracker 1 1.00 0.18 0.65 0.65 0.65 0.65OBC 2.47 2.47 2.47Payload OBC 1 3.60 0.90 2.30 2.30 2.30 2.30Spacecraft OBC 1 0.90 0.01 0.17 0.17 0.17 0.17COM 6.92 4.46 17.67S-Band Ant (Tx) 1 10.70 0.20 10.70 0.00 0.00 10.70S-Band Ant (Rx) 1 0.80 0.20 0.80 0.20 0.20 0.00S-Band Trans. 1 4.13 3.88 4.13 3.88 3.88 4.13UHF Ant (Tx) 1 0.17 0.17 0.17 0.00 0.00 0.17UHF Ant (Rx) 1 0.17 0.17 0.17 0.17 0.17 0.00UHF Trans. 1 3.30 0.18 2.64 2.64 0.18 2.64Dock CM3 2 0.02 0.00 0.02 0.02 0.02 0.02EPS 0.63 0.63 0.63ACU 1 1.70 0.30 0.30 0.30 0.30 0.30PDU 2 0.17 0.17 0.17 0.17 0.17 0.17Payload 4.60 9.50 9.50Camera 1 5.00 0.10 5.00 0.10 5.00 5.00Pumping 1 0.50 0.01 0.50 0.50 0.50 0.50Sensors 1 1.00 0.01 1.00 1.00 1.00 1.00Heating 1 2.00 0.01 2.00 2.00 2.00 2.00Light 1 1.00 0.01 1.00 1.00 1.00 1.00TCS 10.00 10.00 10.00Batt 1 2.00 0.00 6.00 6.00 6.00 6.00Others 4 1.00 0.00 1.00 1.00 1.00 1.00Total 27.88 31.53 44.74Total w/unit&sys margin 33.73 38.15 54.13

Table 4.7: Power Budget and Power Operational Modes.

• Tlight : Time of illumination• PEOL: Power generated at the End Of the Life (EOL)

The energy consumed is determined as follows:

Econs =PnomTnom

ηdist+

(PEOLTcom +

(Pcom−PEOLηdist)Tcom

ηbatηdist

)(4.9)

With

Tnom = Tlight −Tcom (4.10)

• Tcom: Time in nominal+communications mode.• Tnom: Time in nominal mode.• Pnom: Power consumed in nominal mode with margin.• Pcom: Power consumed in nominal+communications mode with margin.

23

Chapter 4. Subsystems: Analysis and Design

• ηbat : Distribution efficiency from the battery to the units.• ηdist : Distribution efficiency from the panels to the units.

Analysing equation 4.9:• First term: represents the energy that is consumed by S/C components in nominal mode.• Second term: represents the energy that is consumed by S/C components in communications mode.

In this mode, the battery is not charged but discharged. This terms has to take into account not onlythe losses from the solar array to the battery (ηdist) but also the losses from the battery to the units(ηbat).

The EOL power is obtained isolating it in the equation. Then the yearly degradation (in percentage) andthe nominal mission duration is considered to obtain the BOL power. Since the mission is really short,this effect is not significant:

PBOL =PEOL

(1− yeardegr)years (4.11)

Finally, solar array inherent degradation, cell efficiency and a conservative value of cosine losses are usedto size the solar array:

Psa =PBOL

Idcosθηcell(4.12)

ncells =Psa

PSUN(4.13)

With• PBOL: Power generated at the Beginning Of the Life• θ : Sun Incident angle• ηcell: Cell efficiency• PSUN : 1366 W/m2

Despite the S/C will be always sun-pointing, 5 deg of cosine losses are considered to be conservative.Cell efficiency and yearly degradation is taken from component datasheet (see trade-off in Section 4.7.4),inherent degradation is taken from literature [22] and experience. Finally, the total number of cells isobtained dividing the obtained area by the cell size (COTS). Obviously this value is rounded up to obtainan integer number. Table 4.8 summarises the values used and obtained in this process.

ncells =Asa

Acell(4.14)

As it was above-mentioned, the battery provides the extra power required for nominal+communicationsmode. This is computed taking the power required for such mode and subtracting the power generationcapacity of the solar array (taking into account all efficiencies and degradation). Besides, the batteryefficiency is considered. The number of cycles for this short mission (90 days times average 25 G/Scontacts per day) does not reach the minimum ones for using the Depth of Discharge (DOD) vs cycleschart of [22]. Hence, theoretically, DOD could be 100%. Nevertheless, to be conservative, a 60% DOD isselected. Table 4.9 shows a summary of the battery sizing.

Ebat [Wh] =(Pcom−PEOLηdist)Tcom

ηbatDOD=

∆PbatTcom

ηbatDOD(4.15)

24

4.7 Electrical Power Subsystem (EPS)

SOLAR ARRAY SIZINGTlight 24 h PBOL req. 46.26 WTcom 1.25 h Inherent deg. 85%Pnom 38.15 W ηcell 32%Pcom 54.13 W Solar Power 1366 W/m2

∆Pbat 14.43 W Solar Array Area 0.1249 m2

ηdist 85% Solar Array Area 1249 cm2

ηbat 85% Cell Area 30.18 cm2PEOL req. 46.03 W Req. cells 41.38 cells numYearly deg. 2% Power per cell 1.12 WMission duration 0.25 years Req. cells round 42 cells numSun Angle 5 deg P_BOL 42cells 46.95 WCosine loss 0.996 P_EOL 42cells 46.72 W

Table 4.8: Solar Array Sizing

BATTERY SIZING∆Pbat 14.43 WTcom 1.25 hηbat 85%DOD 60%Battery Energy 35.4 Wh

Table 4.9: Battery Sizing

4.7.4 COTS Trade-offIn this section, a COTS market research is performed. Several options for each component are presentedand a trade-off according to different aspects with their respective importance/weigh (indicated in eachtable) is carried out. Refer to Section 4.5.2 to see the explanation of the trade-off scores computation.

Solar ArraysMMA Design e-HAWK solar arrays are selected. They offer complete customization (final solution is onefixed panel and two double-deployable ones for a total of 42 cells, see Figure 4.10. Summarizing, 6 serieslines of 7 cells each one) and the highest efficiency cells (up to 32% from Azurspace). However, one ofthe key factors is the number of cells per total panel area that e-HAWK panels have in comparisson withtheir competitors, minimising the panel size.

COTS Eff. Mass [gr] Customization Available Info Configuration TOTALWeight 5 2 4 1 3 Max=45

GOMspaceNanoPower

30% 1500 Limited Complete Moderate 25

MMA Design e-HAWK

32% 1200 Superior Complete Superior 45

ISIS Space 30% 1300 Moderate Limited Limited 26

Table 4.10: Solar Array Trade-off.

BatteryCapacity is not a restricting factor for any of the considered suitable batteries. Hence, other criteria likemass or available information from manufacturer (i.e to ensure compatibility) has been taken into account.

25

Chapter 4. Subsystems: Analysis and Design

Finally, GOMspace BP4 is selected.

COTS Capacity [Wh] Mass [gr] Available Info TOTALWeight 1 2 3 Max=18

GOMspace BP4 38,5 270 Complete 17ISIS iEPS 45 310 Limited 10ClydeSpace 40 335 Limited 7

Table 4.11: Battery Trade-off.

PMADThe PMAD shall match the required number of input/output ports (compatibility, see Figure 4.6). Onceagain, datasheet information available is also considered to ensure that compatibility. GOMspace’s PMADis selected.

COTS Compatibility Mass [gr] Available Info TOTALWeight 3 1 2 Max =18

GOMspace P60 Confirmed 248 Complete 18EnduroSat Not confirmed 292 Moderate 9

Table 4.12: PMAD Trade-off.

4.7.5 S/C connection diagramThe PMAD was selected taking into account the operation voltages stated in datasheets of the rest of theCOTS. Figure 4.6 shows the connection diagram from the PMAD to the finally selected units (whosetrade-offs are performed in their respective sections). The selected GOMspace P60 system includes:• 1 x ACU: fitted with 6 input ports that is the exact number required to integrate the six solar array

lines that are present in the satellite.• 2 x PDU: each of them with 9 output ports. The voltages of these ports can be selected between

3.3/5 or 8 V. Besides, one of these ports can be customized to have other voltage (12/18 or 24 V).

Heaters

S-Band TR

OBC Z-7000

PL Heaters

PL Camara

PL Others

UHF TR

OBC A3200

StarTracker ST-200

4xRWs Magnetorquer

3xSun-Sensors

S-Band Antenna

UHF Antenna

PDU1 PDU2ACU

PMAD

5V

3.3V

3.3V3.3V

3.3V

8V

3.3V

3.3V

5V

5V

3.3V

3.3V3.3V

5V

5V

3.3V

5V

5V

*Note: None of the lines should exceed 2 A current.

Figure 4.6: S/C Components Connection Diagram.

26

4.8 Attitude, Determination and Control Subsystem (ADCS)

4.8 Attitude, Determination and Control Subsystem (ADCS)

The attitude determination and control system (ADCS) is responsible for two main tasks: the attitudedetermination, i.e. the dynamics, kinematics and the environment related to the S/C; and the attitudecontrol of the system, i.e. the set of actions performed by the S/C to maintain a correct attitude state giventhe estimated attitude.

Figure 4.7: Spacecraft nominal attitude in orbit.

According to [22] in order to design the ADCS, several steps need to be taken. The control modes andsystem-level requirements have been already defined in Section 2.4. In this section, the calculation of theenvironmental disturbances, the type of actuators to control the attitude and the sizing of these componentsis developed.

4.8.1 Disturbance sizing

Once the system-level requirements and the control modes have been defined, it is necessary to determinethe environmental disturbances that will affect the S/C attitude during its operation and, therefore,constrains the selection and sizing of the ADCS components. There are four main disturbance torquesthat will affect the attitude of the S/C [22]: gravity-gradient torque, solar radiation pressure, aerodynamictorque and magnetic torque. This disturbance will depend on the mission parameters selected in Section2.1, more specifically, it is necessary to focus on the orbit geometry and payload parameters shown inSection 4.1. Given these parameters and following the expressions provided in [22], the disturbances canbe sized.

Gravity-gradient torque

The gravity-gradient torque occurs when there is a misalignment of the principal axis with respect to thelocal vertical. If the angle between the principal axis and the local vertical increases, this torque will belarge. For this particular case, the local vertical of the S/C is coincident with the radius of the Earth, i.e.the angle deviation is between the minimum principal axis and nadir. According to the requirements (seeTable A.1) the S/C must point to the Sun, therefore the misalignment will vary during the operation of theS/C. Thus, the value provided in the following equation corresponds to the worst possible case, i.e. whenthe angle is maximum. The gravity-gradient torque is computed as follows:

27

Chapter 4. Subsystems: Analysis and Design

Tg =3µ

r3p|Imax− Imin|sin(2θ)→ Tg ' 2.57 ·10−7 N·m (4.16)

• Tg. Gravity-gradient torque around Y-axis.• µ = GM = 3.986 ·1014 m3/s2. Standard gravitational parameter.• rp = REarth +h = 6378 ·103 +500 ·103 = 6878 ·103 m. Periapsis radius.• Imax ≡ Ixx = 0.168 kg·m2. Maximum moment of inertia (see Table 4.24).• Imin ≡ Izz = 0.090 kg·m2. Minimum moment of inertia (see Table 4.24).• θ = 31◦. Maximum angular deviation between the longitudinal axis and the nadir-zenit direction.

As stated before, the worst-case possible is taken for the computation. This case occurs taking intoaccount the inclination (' 7.5◦) and the ecliptic plane of the orbit (' 23.5◦).

Solar radiation pressureThe solar radiation pressure (SRP) torque is produced by the incident radiation of the sunlight in theS/C. It depends mainly on the distance between the Sun and the satellite, being more important in GEOthan in LEO. It also depends on the projected area with respect to the Sun which in this case needs to beconsidered due to the sun-synchronous orbit and sun-pointing requirement. The reflection coefficient ofthe surface pointing the sun, i.e. solar panels, and the centre of gravity need also to be taken into account.SRP torque is given by the following equation:

Tsp =Φ

cAsp(1+q)cos(i)‖csp− cg‖→ Tsp ' 1.73 ·10−7 N·m (4.17)

• Tsp. Solar radiation pressure torque.• Φ = 1412 W/m2. Solar flux. The maximum value for the solar irradiance flux is taken.• Asp = 0.18 m2. Projected area in Sun direction. Due to the sun-pointing requirement the maximum

area with the solar panels deployed is considered.• q = 0.7. Reflectivity coefficient. This value has been selected taking into account the literature

[10, 19] and the parameters of the solar panels showed in Table 4.8.• i = 0◦. Sun light incident angle. This value is set taking into account the sun-pointing requirement.• ‖csp− cg‖ = 0.1 m. Distance between the centre of pressure, csp, and the centre of gravity, cg.

A conservative value is chosen given the centre of gravity (see Table 4.24), the geometry of theCubeSat, the sun-pointing requirement and the literature [22].

Atmospheric drag or aerodynamic torqueTaking into account the orbit geometry chosen in the predesign mission (see Section 4.1), the atmosphericdrag which induces an aerodynamic torque needs to be determined. This torque affects the S/C whichoperates in LEO (mainly below 400 km) because the atmospheric pressure has an impact in the areaexposed to the orbital direction of the satellite if the centre of pressure is misaligned with respect to thecentre of mass. This effect is also higher with higher orbital velocities and larger exposed areas. Thus, theaerodynamic torque will be determined by:

Ta =12

ρatm,pCDAaV 2p ‖ca− cg‖→ Ta ' 4.54 ·10−7 N·m (4.18)

• Ta. Aerodynamic torque.• ρatm,p = 1.74 ·10−13 kg/m3. Atmospheric density at the periapsis. This value is taken following the

MSIS-E-90 Atmosphere Model by NASA1.• CD = 2.5. Drag coefficient. This value is selected following the literature [4].• Aa = 0.18. projected area in the velocity direction. The maximum area with the solar panels

deployed is considered.

1https://ccmc.gsfc.nasa.gov/modelweb/models/msis_vitmo.php

28

4.8 Attitude, Determination and Control Subsystem (ADCS)

• Vp =√

µ

a =√

µ

REarth+h = 7612 m/s. S/C velocity at periapsis. Note that the orbit is circular (seeSection 4.1).• ‖ca− cg‖ = 0.2 m. Distance between the centre of pressure, ca, and the centre of gravity, cg. A

conservative value is chosen given the centre of gravity (see Table 4.24), the geometry of theCubeSat and the literature [22].

Magnetic torque

The magnetic torque occurs due to the interaction between the Earth’s magnetic field and the CubeSatmagnetic field usually called residual magnetic dipole of the S/C. The magnetic torque will depend onthis residual magnetic dipole, which for nanosatellite is of the order of 10−2 [13, 21]. This value is alsodependent on the distance with respect to the Earth, being more important for LEO than for GEO sincethe Earth’s magnetic field is stronger. By definition:

Tm = D ·B≈ DmλMm

r3p→ Tm ' 1.47 ·10−6 N·m (4.19)

• Tm. Magnetic torque.• Dm = 0.03 A·m2. Vehicle residual magnetic dipole. There are different models for the determination

of the S/C residual magnetic dipole for nanosatellites [8][13][21]. The value selected if based onthe literature.• Mm = 7.96 ·1015 T·m3. Earth magnetic moment of the central body [22].• rp = REarth +h = 6378 ·103 +500 ·103 = 6878 ·103 m. Periapsis radius.• λ = 2. Function of the magnetic latitude [22]. Due to the orbit geometry (see Section 4.1) the

function value at the poles is chosen.

Other disturbances

There are other small disturbances and uncertainties that may affect the performance of the ADCS whichneed to be accounted for. In general terms, when sizing the ADCS components, a safety margin is addedto the calculations. For this particular design in which there are no thrusters, these disturbances can besummarised as follows:• Uncertainty in centre of gravity. For this case the centre of gravity is known (see Table 4.24).

However, the CoG may change due to changes in the payload or in the structure once the S/C is inspace. As shown in the previous paragraphs, this may affect the SRP torque and the aerodynamictorque. As aforementioned, a conservative value for the distances between the CoG and the centreor pressure (solar and aerodynamic) has been considered in order to cover this disturbance.• Thermal shocks on flexible appendages. The thermal shocks may occur during eclipses affecting

the performance of gravity-gradient systems and other components. This effect is minimised sincethe launch window avoids eclipse phases along the science operation (see Section 4.1).• Reaction wheel friction and electromotive force. This disturbance is caused by the reaction

wheels in which any kind of friction reduces the speed of the wheel. This affects the performanceof the reaction wheel and needs to be accounted during the sizing of the ADCS components.• Rotating machinery (pumps, filter wheels), liquid slosh and others. Due to the uncertain

behaviour of the payload, it can produce small disturbances that need also to be taken into accountwhen determining and sizing the ADCS components.

Final results

After the computation of the main disturbances, it can be concluded that the magnetic torque is themost determinant disturbance for the ADCS components sizing, followed by the contributions of theaerodynamic torque, the gravity-gradient toque and, finally, the solar radiation pressure. Table 4.13summarises the results and computes the total external torque suffered by the satellite. A safety margin of20% is applied to the calculation to be conservative.

29

Chapter 4. Subsystems: Analysis and Design

Magnetic Torque 1.47 ·10−6 N·mAerodynamic Torque 4.54 ·10−7 N·mGravity-gradient Torque 2.57 ·10−7 N·mSolar Radiation Pressure (SRP) Torque 1.73 ·10−7 N·m

Total Torque 2.31 ·10−6 N·mTotal Torque + Safety Margin (20%) 2.78 ·10−6 N·m

Table 4.13: Disturbance sizing for the ADCS.

4.8.2 Selection and Sizing of ADCS componentsOnce the system-level requirements and control modes are exposed, and the main environmental distur-bances for the operation of the satellite are defined, it is necessary to define, select and size the componentsthat will conform the ADCS.

The first step consists on determining whether the attitude control of the ADCS will be active control,passive control or both. In order to select they type of attitude control it is necessary to look at therequirements (see Table A.1). More precisely, the high accuracy pointing system-level requirements(NS-ADC-01) discards the use of only passive control components. Based on an iterative process in whichseveral components have been tested to fulfil these requirements and based on the tips provided in theliterature [22] and [5], active control components have been selected to shape the ADCS.

Furthermore, it is necessary to determine which techniques or components among the active control willbe selected. Again, taking into account the system-level requirements of pointing accuracy (NS-ADC-01,NS-ADC-02) a 3-axis control technique based on four reaction wheels and a magnetorquer is proposed.Thus, in terms of sensors, the use of a star tracker and three sun sensors are needed (a magnetometer isalso included in the on-board computer). This kind of attitude control components are nowadays used andalthough they are more expensive they provide an accurate, stable and reliable solution. The 3-axis controlcomponents, i.e. reaction wheels and magnetorquer, allow to counteract the environmental disturbancesand preserve the orientation of the CubeSat within the requirements proposed. Below, the main parametersthat size the ADCS components are developed.

Reaction wheelsThe reaction wheels are active control components that can generate a counteracting torque through theirrotation, having one axis of control for each wheel. The torque will depend on the size of the wheel andits spinning velocity. As a drawback, reaction wheels need time for de-saturation and their control lawsare challenging.

In order to size the reaction wheels, there are two main quantities that need to be computed. The firstof them is the maximum torque that the reaction wheel is capable of delivering. Obviously, this torquemust be higher than the maximum total torque produced by the environmental disturbances (with thesafety margin applied for small disturbances). Furthermore, the slew torque (and slew rate) needs to becomputed, since it is the torque related to keeping the S/C with the sufficient accuracy to perform thecontrol modes and maintaining the attitude in an accurate and stable state. Thus,

TRW = max{TRW,D,TRW,S} (4.20)

where TRW,D is the total disturbance torque with the safety margin calculated in the previous section. Theslew torque TRW,S is computed considering that the S/C rotates 360 degrees in an orbital period. Thus theslew torque and slew rate are defined by:

30

4.8 Attitude, Determination and Control Subsystem (ADCS)

TRW,S = 4θπ

180Imax

t2s' 1.31 ·10−7 N·m (4.21)

θ̇S =θ

ts= 0.06 deg/s = 0.0011 rad/s≈ 0.01 rpm (4.22)

• θ = 360◦. Maximum angle of rotation.• ts = 2π

√rpµ= 5676 s. Minimum duration of the slew manoeuvre.

• Imax ≡ Ixx = 0.16787 kg·m2. Inertia axis of the slew manoeuvre. The maximum axis is taken forconservative analysis.

The slew rate is low and it will be assumable by the reaction wheels according to [22]. Finally, themaximum torque that needs to be provided by the reaction wheel is:

TRW = max{TRW,D,TRW,S}= TRW,D = 2.78 ·10−6 N·m

However, the slew torque required will be also taken into account when performing the selection of thecomponents in the next section.

The second parameter that needs to be studied is the momentum storage in the reaction wheels. This is thetotal angular momentum that the reaction wheel needs to withstand and is given by:

HRW = TD ·P · fdisturbance ' 2.78 ·10−3 N·m·s (4.23)

• TD = 2.78 ·10−6 N·m. Total disturbance torque with safety margin for being conservative.• P = 2π

√rpµ= 5676 s. S/C orbital period.

• fdisturbance = 0.177. Value based on a function that determines when the maximum momentum isaccumulated and will depend on the orbit geometry and the contribution of the different disturbancetorques [22].

Apart from this two main parameters, there is another torque that needs to be analysed and that affects thelaunch and early orbit phase mainly. Once the S/C is released from the launcher, it suffers a torque thatproduces a nutation in the S/C. The nutation is higher for higher spin-rate tumbling. In order to counteractthis effect, the actuators need to generate a detumbling effect. In order to size this disturbance, there areseveral works [17, 6] that have developed methods for the detumbling cases in which the angular rate ofstudy is between 0 and 10 deg/s (∼ 1.7 rpm). According to the requirements (NS-SYS-07 and NS-ADC-05,see Table A.1) the maximum and limiting rotation speed is higher than the one provided. Nevertheless,the maximum time allowed for the angular rate of the requirements is used for the calculations, i.e. 10seconds. Thus:

Tdetumbling =dHdt≈ Imaxωrelease

tdetumbling≈ 2.93 ·10−3 N·m (4.24)

Note that this is a critical case and since, for this case, the reaction wheels are operating at a lower angularrate than the limiting one, the detumbling time could be higher, therefore, reducing the torque.

MagnetorquersTo add control redundancy, a magnetic torquer has been implemented. This device is another type of activecontrol actuator, which consists on a magnetic coil that generates a magnetic dipole. Since the maximumdisturbance torque for this particular case is the magnetic moment, the use of a magnetorquer is a greatsolution as magnetic torquers allow to counteract small disturbances and those related to the magneticfields and residual magnetic dipole of the S/C. As said before, it will be used as a system redundancy anddesaturation for reaction wheels, as the time to compensate the disturbance effects is much larger than forreaction wheels. Its main characteristic is the magnetic dipole that it is able to produced and is given by:

Dm,MT =TB' 6.06 ·10−2 A· m2 (4.25)

31

Chapter 4. Subsystems: Analysis and Design

• T ≡ TD = 2.78 ·10−6 N·m. Total disturbance torque with safety margin for being conservative.• B = 4.58 · 10−5 T. Earth’s magnetic field at 500 km. This value has been obtained using the

International Geomagnetic Reference Field2 model at the critical case, i.e. at the poles.

Sensors

The sensors are key components in the ADCS. They allow to calculate the attitude determination of theS/C and they will be selected according to the requirements. Again, within the requirements, the mostdetermining ones are those related to the pointing accuracy. In order to provide a robust solution, a startracker and three sun sensors are used. Moreover, the addition of a magnetometer is necessary in orderto take into account the presence of the magnetorquer. A brief summary of the main properties andcharacteristics of these devices is exposed in the following lines.

• Star tracker. This device is composed by optics and a focal plane with an electron detector, togetherwith detection electronics. The star tracker detects on its sensing device the light emitted by stars inthe Field of View (FoV), compares the position of the stars in the FoV with the position of stars in astar-catalogue and then, computes the orientation of the S/C. It provides as measurement directly aquaternion with respect to the celestial frame. These devices have an update rate between 0.5 and10 Hz and have an accuracy of the order of arc-secs with a FoV of ±6◦ [7]. Apart from its powerconsumption, mass and accuracy, it is necessary to take into consideration the Noise EquivalentAngle (NEA) which is the angular white noise error on the STR. Note that the star tracker provideslow-frequency absolute measurements, not subject to drifts.• Sun sensors. These devices detect the line-of-sight of the Sun to provide an approximate unit

vector with respect to the body reference frame that points towards the Sun. Their accuracy is ofthe order of 10−2 degrees with a typical FoV of ±30◦ [7]. Apart from the redundancy that thesedevices provide they are a perfect option for sun-synchronous orbit like the one chosen and alsoprovide a higher FoV to avoid errors in the attitude calculation.• Magnetometer. As said before, the use of a magnetorquer makes necessary to include a magne-

tometer. This device provides a vector with the Earth’s magnetic field obtained with respect to itsfixed reference frame. For its calibration it is important to take into account the residual magneticdipole generated by the magnetic torque. A good practice is to locate the magnetometer as far aspossible from the magnetic torquer in the geometry of the S/C. They are only valid for LEO (like inthis particular case) and provide an accuracy of the order of 1◦−5◦ [7]. The magnetometer is alreadyimplemented in the on-board computer and no trade-off will be performed to this component.

Note that the S/C tracking is made from the ground station. The methodology for the orbit determinationapplied in the G/S is a two-way tracking, i.e. signal generated at a G/S is transmitted to the S/C andretransmitted to G/S. Once the G/S send the position and velocity to the S/C then the guidance, navigationand control (GNC) software needs to compute the needed algorithms and use the feedback of the G/Sto propagate the resultant trajectory. The result will be send to the G/S and to the actuators in order toperform the proper manoeuvres to control the attitude. For S/C at LEO Doppler data is used in which theaccuracy of orbit determination is of the order of 10−2 mm/s.

4.8.3 COTS trade-off

The last step is to perform a trade-off analysis to select the components that will shape the ADCS. Themethodology for the analysis has been aforementioned in Section 4.5.2. The trade-off has been done forthe different components of the ADCS: reaction wheels (see Table 4.14), magnetorquer (see Table 4.15),star tracker (see Table 4.16) and sun sensors (see Table 4.17).

2https://www.ngdc.noaa.gov/IAGA/vmod/igrf.html

32

4.8 Attitude, Determination and Control Subsystem (ADCS)

COTS Mass (g) Power(mW)

M stor.(mNms)

Max. T(mNm)

Controlacc. (rpm)

Availableinfo Total

Weight 1 2 2 2 2 1 Max=30

HyperionRW210

32 600 ±6 ±0.1 0.5 Complete 24

MAI-400 90 Unknown ±9.35 ±0.635 Unknown Limited 18CS CWSmall

60 600 ±1.7 ±0.23 5 Complete 21

Table 4.14: Trade-off study for reaction wheels.

COTS Mass (g) Power (mW) Mag. Dipole (A/m2) Available info TotalWeight 1 2 2 1 Max = 18

ISIS - iMTQ 196 1200 0.2 Complete 15GOMspaceGST-600

156 Unknown 0.31 Moderate 12

GOMspaceZ-axis Int

106 Unknown 0.139 Moderate 9

MAI-400 694 2050 0.15 Complete 12

Table 4.15: Trade-off study for magnetorquer

COTS Mass (g) Power(mW)

Slew rate(deg/s)

3σ (point/roll)(arcsec)

Up. rate(Hz)

Availableinfo Total

Weight 1 2 2 2 1 1 Max = 27

HyperionST200

42 600 0.3-0.6 30/200 5 Complete 24

BSTST200

40 650 Unknown 10/Unknown 5 Limited 17

Space MicroMIST

520 3000 1-5 6/40 10 Limited 16

Table 4.16: Trade-off study for star trackers.

COTS Mass (g) FoV (deg) 3σ (deg) Up.rate (Hz) Available info TotalWeight 1 2 2 1 1 Max = 21

SolarMEMSnanoSSOC-D60

6.2 120 0.3±0.1 50 Complete 19

NewSpaceNFSS-411

5 114 0.5 10 Moderate 14

NewSpace - 2.2 120 2 Unknown Limited 13

Table 4.17: Trade-off study for sun sensors.

According to the results of the trade-off analysis, four reaction wheels RW210 have been selected inorder to counteract the potential failure of one of them (they will be set following a pyramidal structure).One magnetorquer iMTQ has been selected to provide a robust attitude control solution. For sensors,a ST200 star tracker is chosen to achieve an outstanding attitude determination accuracy and will be

33

Chapter 4. Subsystems: Analysis and Design

complemented with three sun sensors nanoSSOC-D60. The performance of each of the components fulfilsthe requirements established in Table A.1 and also are able to provide a consistent solution with enoughsafety margins for the sizing performed in the previous section.

4.9 Mechanical design and structure

The purpose of the structure of the spacecraft is to withstand the external stresses that it will sufferduring its operational time, i.e. from its launcher phase to disposal. Apart from the external stresses, agood structure should resist impact from space debris in order to preserve its internal components in asafe state. Moreover, due to the environmental radiation levels at which the spacecraft is operating, thestructure is fundamental in order to avoid this radiation, produced by several factors, affecting the internalcomponents.

4.9.1 Mechanical DesignAssumptions

The process of designing the mechanical structure of the S/C is challenging and several factors need tobe determined. First of all, it is necessary to take into account that characterising the external stressesthat will be suffered by the spacecraft, is a very challenging task. FEM models are used to shape the finalmechanical design. The load uncertainty, the different types of materials, the environmental uncertainty,etc. are some of the reason for this. In order to be conservative with the analysis, several assumptionsneed to be made for this preliminary design:

• Only the material of the primary structure is considered, i.e. the external and load-carrying.• After the calculations, a safety margin is applied.• The maximum allowable load for the material selected is considered.• For the preliminary design, the stresses exerted during the launching phase are studied. These

stresses are considered to be the most critical during the operation of the satellite.

Material selection

The second step is to determine the material in which the main structure will be manufactured. In thispreliminary design, Aluminium 6061 has been selected. This type of material is mainly used for 6UCubeSats 3 due to its good performance in terms of high strength, low weight, ease of manufacturing,ductility and low density.

Launcher frequencies

As said before, one of the critical phases for the structure is the launcher phase. The structure needs towithstand the stresses produced in this phase. The launcher is selected by the costumer and its naturalfrequencies will determine the thickness of the structure. According to the literature [11][14][20], agood practice is to first size the thickness of the S/C for a range of frequencies. Note that the higher thefrequency, the higher the thickness so, for the preliminary design, the most critical frequency is consideredfor the calculation of the thickness. Since the structure needs to withstand natural frequencies larger thanthe ones of the launcher, if the thickness of the primary structure is larger than the one obtained, therequirements will be fulfilled.

There are two types of natural frequencies that are considered: axial and lateral. Its formulation and thederivation of the thickness is as follows:

fax =1

2π

√EAmL

=1

2π

√EπrtmL→ t = (2π fax)

2 mLEπr

= 1.5×10−4 m = 0.15 mm (4.26)

3https://sst-soa.arc.nasa.gov/06-structures-materials-and-mechanisms

34

4.9 Mechanical design and structure

flat =1

2π

√3EImL

=1

2π

√3Eπr3t

mL3 → t = (2π flat)2 mL3

3Eπr3 = 1.68×10−5 m = 0.017 mm (4.27)

• fax = 300 Hz. Axial natural launcher frequency. This value has been selected from a range proposedin several references [1][23].• flat = 300 Hz. Lateral natural launcher frequency. This value has been selected from a range

proposed in several references [1][23].• E = 68.9 GPa. Young’s modulus of Aluminium 6061.• r = 0.15 m. Radial distance. Since the 6U CubeSat is a cube, the maximum lateral distance from

the centre is considered.• t. Thickness in meters.• m = 15.75 kg. Mass of the spacecraft (see Table 4.25).• L = 0.087 m. Distance from the bottom to the centre of mass.

Thus, the limit thickness is the one obtained from axial launcher natural frequency. To be more con-servative, as stated before, a safety margin of 20% is added. The minimum thickness then shouldbe:

tmin = tax ·SM = 0.18 mm

The thickness of the preliminary design is 1.5 mm (2.6 mm if the thickness value found for shielding inthe next subsection is added). Therefore, the requirement for launcher frequencies is fulfilled (NS-STR-02,see Table A.1).

Launcher loadsAnother structural constraint is given by the loads exerted during launcher phase. These loads will exert aforce that need to be supported by the structure. Again, the primary structure made of Aluminium 6061is considered for the calculation. A simplified formulation of the axial and lateral forces exerted duringlauncher phase is given by:

Paxial = mgLFax ' 1205 N (4.28)

Plat = mgLFlat ' 464 N (4.29)

• Paxial,Plat . Axial and lateral forces.• m = 15.75 kg. Mass of the spacecraft (see Table 4.25).• g = 9.81 m/s2. Gravity constant.• LFaxial = 7.8, LFax = 3. Axial and lateral load factors. These values are taken according to different

resources for most critical values [1, 23] (considering static and dynamic loads load factors).As expected the axial force is larger. However, the lateral force exerts a momentum that will increase thetotal axial force and needs to be considered.

Peq = Paxial +2Mr

= Paxial +2Plat

rL' 1743 N (4.30)

• Peq. Total or equivalent axial force.• r = 0.15 m. Radial distance. Since the 6U CubeSat is a cube, the maximum lateral distance from

the centre is considered.• L = 0.087 m. Distance from the bottom to the centre of mass.

Finally, it is necessary to check if the ultimate and yield conditions of the material can withstand theseloads. In order to be conservative, a safety factor will multiply the results. Thus:

Peq fu

A=

Peq fu

2πrt≈ 0.89MPa < σu = 150 MPa (4.31)

Peq ftA

=Peq ft2πrt

≈ 0.89MPa < σt = 83 MPa (4.32)

35

Chapter 4. Subsystems: Analysis and Design

• ft , fu. Safety factors for ultimate tensile strength and yield stress. Normal values are 1.25 and 1.1respectively, however, it has been selected as value a safety factor of 1.25 not only to asses theperformance but also for later qualification assessment.• σu, σt . Ultimate yield strength and yield stress. Data obtained by the Aluminium 6061 standards4

[2].• t = 2.6 mm. Total thickness including shielding.

As expected, the Aluminium 6061 can withstand the load exerted during launcher phase.

4.9.2 S/C ShieldingDuring its life time, the CubeSat will be immersed in an hostile radiation environment where it has tosurvive and operate successfully. Particle impacts can cause significant effects on electronic devices andmaterials. One way to mitigate this effect is to protect the CubeSat with an external layer of Aluminiumshielding. In order to size the required shielding, the Stopping Power is studied [18]. This concept refersto the average energy lost by a charged particle per unit distance while interacting with matter. The Betheformula is an approximation of the Stopping Power for a heavy charged particle and provides the result inkeV/µm:

S =5.08×10−32z2n

β 2

[ln(

1.02×106β 2

I (1−β 2)−β

2)]

(4.33)

Where:• z = q/e = 1 and e is the electron charge• β = v/c, v is the velocity of the particle and c is the speed of light.• n is the electron density of the absorber and only depends on the material (NA is the Avogadro

number, Mu ≈ 1g/mol, ρ is the density, Z is the atomic number and A is the atomic mass number):

n =NAZρ

AMu(4.34)

• I = 16×Z0.9 is the mean excitation energy in eV .

According to ECSS-E-ST-10-12, E = 10MeV protons are considered to infer the velocity of the particle(through kinetic energy). The stopping material is Aluminium 6061 whose density is 2700kg/m3, Z = 13and A = 27. Finally, the ratio between the particle energy and the computed stopping power provides anestimation of the required aluminium shielding thickness (R). Table 4.18 summarises the obtained results.

R =ES

(4.35)

Stopping Power 9.18 keV/µmEnergy 10 MeVShielding thickness 1.1 mmShielding mass 0.352 kg

Table 4.18: Shielding estimation

In addition, according to the mission statement, the PL has its own 0.8 mm aluminium shielding. On topof the computed shielding mass, an uncertainty margin is taken in the mass budget. In any case, this is justa rough estimation and TID analysis, SEE assessment, etc. shall be performed in detailed design phases inorder to check whether the COTS radiation tolerance (normally given in krad) is surpassed or not.

4https://www.astm.org/Standards/B209.htm

36

4.10 Thermal Control System (TCS)

4.10 Thermal Control System (TCS)

The thermal control subsystem is devoted to maintaining the operating temperature of every other sub-system. In this context, it is possible to define the requirements that a properly designed thermal controlsubsystem must fulfil. Such requirements come to light when the temperature of each unit in the CubeSatis constrained (ECSS-E-ST-31C).

Each component has a predefined qualification temperature range, in which it has been demonstrated thatoperation is safe and performance is appropriate (included in components datasheet). However, accordingto ESA standards (see Figure 4.8), a qualification margin shall be applied to account for more severeconditions during the mission than those considered in the study (assumed as 5oC according to industrystandards). Besides, an acceptance margin devoted to minimise the impact of unpredictable events overthe complete system is required (also assumed as 5oC considering the industry standards). Finally, anothermargin has to be considered for the uncertainty of the TMM that has been developed. This source of errorare added on top of the calculated temperature ranges.

Figure 4.8: Temperature ranges according to ECSS-E-ST-31C.

Another constraint imposed in this mission is that the TCS shall not include any active TCS system otherthan heaters (like thermal louvres or heat pipes). It should only relay on passive insulation (MLI), thermalderivations (conduction) and simple active heating (resistances). Both a hot and a cold case are studied tocover the extreme conditions.

4.10.1 Thermal Mathematical Model

In this section, the model that has been used to design the TCS in order to comply with the requirementsis detailed. This model starts from the energy conservation equation to provide an energy balance on eachcomponent. According to this idea, equation 4.36 is fulfilled at every node:

mi ·Cpi

δTi

δ t= Qsun,i +Qearth,i +QAlbedo,i +

N

∑j=1, j 6=i

Qrad, j→i +N

∑j=1, j 6=i

Qcond, j→i +Qdiss,i +Qh,i (4.36)

37

Chapter 4. Subsystems: Analysis and Design

Where:

Qsun,i = Ssun ·Ai ·αi ·Fsun,i(t)

Qearth,i = Searth ·αi ·Ai ·Fearth,i(t)

QAlbedo,i = Ssun ·Ai ·αi ·ρearth ·FAlbedo,i(t)

Qrad, j→i = σ · εi, j ·Ai ·Fi, j · (Tj(t)4−Ti(t)4)

Qcond, j→i =Ai, jRi, j· (Tj(t)−Ti(t))

Qdiss,i = Qdissipation(Op. mode(t))

Qh,i = Vi(t)2/Ri

Heat dissipation indicates the power consumption of each element to provide its function, including, inthe case of the PL, the pumping, imaging, and LED lighting, whose load profile has been considered inthe model (2.5 hours ramp up, 1 hour at 1 constant watt, and 2.5 hours ramp down). The operational modehas also been considered, since during communication phases, some units consume different power thanat nominal operation mode. Finally Table 4.19 includes the main variables involved in the equation.

Ssun Solar constant 1366 W/m2

Searth Earth emitted radiation at orbit altitude σT 4rad,earth(

Rearth+hRearth

)2

Ai Node area ∑Alateral

ρearth Earth reflectivity ' 0.2αi Node absorptivity SUN ' 0.6εi, j Cross emissivity εi·ε j

εi+ε j−εi·ε j

Fsun,i Fraction of Sun radiation received cos(φi)

Fearth,i Fraction of Earth radiation receivedView factorsFAlbedo,i Fraction of Sun radiation reflected by the Earth received

Fi, j View factor between nodes i and jQdiss,i Heat dissipated by each node

Heating conditionsQh,i Heating power

Table 4.19: Clarification of the variables included in the thermal balance equation.

Model assumptions

The thermal model used in this study considers the following assumptions and simplifications, whichrelate the mission requirements:

• Each CubeSat component is an isothermal node. Their geometry has been modelled as a paral-lelepiped for most of the nodes.• Nodes can only exchange energy through cross conduction/radiation and exterior radiation. No

possible convection or mass outflow (no propulsion is allowed). Figures F and F.1 in appendix F.2contain the radiation and conduction interconnections chart of all the nodes in the system.• Nodes can only generate thermal energy through dissipation, including energy employed to provide

the services (depending on the node) and the energy employed by the TCS to heat punctually certainnode.• The surfaces of the nodes are grey (excepting MLI) and diffuse.• Study is non-stationary, and involves cyclic variables that may not be in-phase.• The mass, dimensions, and thermal properties of the nodes are obtained from their datasheets or

inferred from their materials.• Simplifications: view factors below 0.01% are neglected, since Montecarlo method is not accurate

enough in such case, and the effects are negligible.

38

4.10 Thermal Control System (TCS)

View factors

The view factor between different nodes represents the fraction of radiation emitted by the first one thatreaches the second. This magnitude is often obtained integrating the solid angle that one surface has as aperspective to the other. A Matlab code has been developed to calculate these integrals using a stochasticmethod (Montecarlo method for integrals) with a considerable accuracy (ε < 1e−4).

On the other hand, the fraction of radiation received from the Earth and the reflected radiation have beencalculated by exact integration. Such factors are dependent on the orbital angles (altitude, attitude andright ascension), what makes them dependent on time. This relationship has been studied and implementedin the TCS analysis.

Equation resolution

For each instant of time, every magnitude is evaluated. Qdiss depends on the power consumption of eachelement regarding the operation point (nominal, communication with ground station, or safe mode). Theheating power is turned on whenever the node temperature approaches the cold limit (design temperaturerange, see Figure 4.8). The equation is integrated numerically using a Runge-Kutta method of 4th order(implemented manually).

Iteration process

The previous model allows to compute the evolution of temperature along time. Different variables of thethermal design can be modified in order to comply with the requirements. Such design variables are:

• MLI absorptivity: allows to increase the total solar power entering to the CubeSat (more effectivein the hot study case).• MLI internal emissivity: allows to transfer thermal energy to the interior of the CubeSat (avoids

reflection to outer space).• MLI external emissivity: allows to dissipate energy to outer space (needs to be controlled in the

cold case).• PL emissivity (if insulated with an MLI): allows to insulate the PL from the temperature oscillations

inside the CubeSat (increases the impact of PL thermal inertia vs. heat flows).• Conduction bridges: allow to control the average temperature point of any node that requires it

(either to increase the temperature or to decrease it).• Installed heater power: allows to correct slight deviations of the temperature below the lower bound

of the design range.

The iterative process consists of adapting the variables progressively in order to comply with the ECSS-E-ST-31C, starting from those which provoke a broader effect on the CubeSat, to those which affect one orfew nodes. Then, the cold case is studied, and, if needed, a wider iteration process will be performed.

Hot and cold case studies

This mission has been designed to avoid eclipses using an SSO LEO. For this reason, the expectationis to only deal with a hot case where the sun strikes directly on the CubeSat. However, depending onthe launching date or if the mission period is extended, there is a possibility to have a 20 minutes Moonpenumbrae eclipse (70% maximum penumbrae). Both the hot and the cold case are studied (see Table4.20).

39

Chapter 4. Subsystems: Analysis and Design

Magnitude Hot case Cold caseGsun ·Fsun,i Constant (1366 W/m2) Variable [0.7 ·Gsun, Gsun]

˙Qdissipated Nominal or communication mode Possibly safe mode˙Qheating Only if needed Only if needed and if possible

FAlbedo,i Same method to compute, but different values due to declination

Table 4.20: Comparison between Hot and Cold cases.

4.10.2 TCS design: iteration and solution

0 20 40 60 80 100 120 140 160 180

Time [min]

260

270

280

290

300

310

320

Te

mp

era

ture

[K

]

Internal nodes temperature

Payload

S-Band Transceiver

UHF Transceiver

SpaceCraft On-board computer

Payload On-board computer

Battery

Power management and distribution

0 20 40 60 80 100 120 140 160 180

Time [min]

282

283

284

285

286

287

288

289

Te

mp

era

ture

[K

]

Payload temperature

Payload

0 20 40 60 80 100 120 140 160 180

Time [min]

284

285

286

287

288

289

290

291

Te

mp

era

ture

[K

]

Payload temperature

Payload

0 20 40 60 80 100 120 140 160 180

Time [min]

279

280

281

282

283

284

285

286

287

288

289

Te

mp

era

ture

[K

]

Payload temperature

Payload

Figure 4.9: Thermal Results Iterations. Top-left 1st It; Top-right 2nd It; Bottom-left 3rd It; Bottom-right4th It.

• Initial approach: the usual values for the emissivities and absorptivities are applied (approximately0.6 excepting the MLI, whose emissivity depends on necessities). Densities and heat capacities areinferred from the materials, average weights and average thermal properties (inferred from COTSdatasheets). The result obtained provides a too cold internal temperature in the CubeSat.• First iteration: MLI radiative properties are adapted to increase average CubeSat temperature.

Also, external emissivity is decreased to avoid heat loss.• Second iteration: the temperature oscillations in the PL are too wide, so it is not possible to correct

directly the average temperature in order to comply with the requirements. It is necessary firstto insulate the PL by conduction (thermal insulator) and radiation (using a MLI that covers itscomplete surface).• Third iteration: the average PL temperature goes over of design range. The solution is to connect

PL by conduction to a cold source to adapt reference oscillation temperature. The selected surface

40

4.10 Thermal Control System (TCS)

is the bottom structure panel.• Fourth iteration: considering the cold case, PL and other components can exceed their lower

temperature limit. For this reason, they are connected to a hot node and PL emissivity is furtherreduced. The result is compliant with the PL requirements. However, a fifth iteration is required forthe rest of the S/C. After that, the temperature ranges are compliant with the design ranges for boththe hot and cold cases.

4.10.3 Requirements ComplianceThis section states the compliance with the temperature ranges that each component needs to operatewithin. Appendix F contains the temperature curves along time of all the components, as well as theiracceptance range (including uncertainty). On the other hand, it also contains figures of this curves in thecold case study.

Results analysisThe study of both hot (Figure F.10 ) and cold cases (Figure F.16) show some conclusions. Firstly, theeffect of the oscillations of the Albedo and Earth radiation factors are the most relevant in the analysis(together with the Moon penumbrae eclipse). The effect of the heat dissipation of the nodes provokes slightdifferences in the temperature range, and the heating power only provides small temperature incrementswhen the temperature decays slowly (only in highly insulated units). On the other hand, the resultsanalysis show that the TCS has a high conditioning number, leading to the idea that more development andinformation is required to provide a fully reliable system. Since slight variations of the model parametersprovoke relevant differences in the TCS output, it is needed to take special care with the PL temperaturecontrol, like a TMM using ESATAN. Appendix F includes all temperature evolution charts of the nominalcase and the most challenging one of the cold case.

TCS parametersThe design meets all the requirements. The final design parameters are stated in Table 4.21:

MLI absorptivity αMLI = 0.6MLI exterior emissivity εMLI,ext = 0.36MLI interior emissivity εMLI,int = 0.6

PL emissivity (MLI insulation) εPL = 0.0005PL conduction insulation Rth,PL ≥ 104K/WPL conduction derivation Rth,PL−node2 ' 0.625K/W

Battery emissivity (colouring insulation) εbat = 0.2Battery conduction insulation Rth,bat ≥ 104K/WBattery conduction derivation Rth,bat−node1 ≤ 5 ·10−3K/W

Battery auxiliary heating Qh,bat = 4WMagnetorquer conduction insulation Rth,magn ≥ 104K/WMagnetorquer conduction derivation Rth,magn−node1 ≤ 0.02K/W

Star tracker conduction insulation Rth,ST ≥ 104K/WStar tracker conduction derivation Rth,st−node1 ≤ 0.02K/W

Table 4.21: Design parameters of the TCS.

Component Temperature RangesTable 4.22 shows the maximum and minimum temperature along time for each node taking into accountmodel uncertainty (see Figure 4.8). Uncertainty of the TMM model is considered as 10oC for most ofthe units, 3oC for such units that include an active thermal control mechanism, 0oC for the PL (since itincludes its own thermal regulation mechanism), and 5oC for the units that show no significant oscillationsalong time (solar panels). The following table shows the design range for each unit in contrast with the

41

Chapter 4. Subsystems: Analysis and Design

predicted maximum and minimum temperatures in both hot and cold case (calculated temperatures +model uncertainty):

Node ID ComponentDesign temp. range Hot case ±ε Cold case ±ε

Min. T Max. T Min. Max. T Min. Max. T1 Front panel -50 80 52.11 73.33 23.76 73.342 Back panel -50 80 -33.23 -10.55 -52.81 -10.553 Right panel -50 80 -14.82 7.69 -36.389 7.694 Left panel -50 80 -22.03 0.73 -42.95 0.735 Top panel -50 80 -22.90 -1.21 -43.62 -1.226 Bottom panel -50 80 -10.72 34.69 -20.67 34.697 Payload 11 15 12.36 15.00 11.15 15.008 S-Band transceiver -30 75 8.26 29.94 -15.51 29.949 UHF transceiver -20 75 9.07 30.77 -14.85 30.78

10 SC-OBC -20 75 11.75 33.42 -12.12 33.4211 PL-OBC -30 75 0.92 22.69 -22.17 22.6912 Battery 10 35 25.76 33.26 13.54 33.2713 PMAD -25 75 10.12 31.82 -13.86 31.8214 Magnetorquer -30 75 2.42 24.04 -19.98 24.0415 Reaction wheels -17 50 6.92 28.62 -16.69 28.6216 Star tracker -17 30 2.37 23.88 -19.29 23.8817 Central solar panel -50 110 97.77 109.99 70.29 109.9918 Right solar panel -50 110 50.05 60.08 22.08 60.0819 Left solar panel -50 110 49.95 59.98 21.99 59.9820 S-Band antenna -30 75 -24.35 28.16 -29.13 28.1621 UHF antenna -30 80 -24.35 28.16 -29.13 28.1622 MLI front panel -110 110 77.11 97.74 46.49 97.7423 MLI Back panel -110 110 -68.52 -46.23 -85.34 -46.2324 MLI Right panel -110 110 -35.66 -14.34 -55.95 -14.3425 MLI Left panel -110 110 -40.08 -18.68 -59.99 -18.6826 MLI Top panel -110 110 -59.59 -38.14 -77.39 -38.1527 MLI Bottom panel -110 110 -7.38 48.12 -11.73 48.12

Table 4.22: Temperatures at each node. Compliance with requirement NS-SYS-05

Conclusion

The preceding analysis has led to a design of a thermal control system that fulfils all the temperaturerequirements. On top of that, it has led to the conclusion that the thermal study needs more preciseinformation and analysis. For this reason, a finite element thermal analysis is proposed (our TMM hasconsidered isothermal nodes instead) to ensure requirements compliance.

4.11 S/C configuration

The internal S/C layout is depicted in Figure 4.11. The L-shaped payload is placed in the bottom part ofthe CubeSat. Then, the majority of the rest of components are placed just above the payload with the ideaof easing the thermal control management and study (the radiation/conduction of the rest of componentscan help to heat up the payload if the thermal engineer determines that it is necessary). Besides, the sunsensors are distributed in panels other than the solar arrays, since they incorporate their own sun sensors.With this, the Sun position will be determined soon after CubeSat deployment.

42

4.11 S/C configuration

The stack and deployed configurations of the CubeSat are shown in Figure 4.10. In stack configurationsolar panels and UHF antenna rods are retracted (launch phase). In deployed configuration, the satelliteis presented during its nominal operation phase and as it can be seen, solar panels and antenna are fullydeployed. In any case, there are four constrains that have driven the complete CubeSat layout:

1. Solar array configuration2. Antennas: Both the UHF and the S-Band antenna are placed in the -Z panel, since the attitude law

will make this panel to face the Earth continuously for G/S contact.3. Star Tracker: Placed in the top part of the S/C. This size will never be blinded by light sources

like the Earth or the Sun and will always face the stars. In addition, this position allows for a startracker baffle addition if future studies determines it (out of the scope of this PDR, and according toECSS-E-ST-60-20C Rev.2) since there is about 5 cm of pace available below it.

4. Magnetorquer: positioned far away from any computer which can be affected by electromagneticinterference. This surely will help in electromagnetic compatibility studies to be performedaccording to handbook ECSS-E-HB-20-07A (out of the scope of this PDR).

Finally the CoG and inertia properties of the CubeSat are showed in Tables 4.23 and 4.24.

Mass[kg]

COG Inertia matrix at COGx[mm]

y[mm]

z[mm]

lxx[kg.m2]

lxy[kg.m2]

lxz[kg.m2]

lyy[kg.m2]

lyz[kg.m2]

lzz[kg.m2]

11.44 6.11 1.86 87.32 0.15412 0.0031 0.0028 0.0963 -0.0072 0.0821

Table 4.23: S/C Mass, CoG and Inertia without margins in Deployed config.

Mass[kg]

COG w/ marg. Inertia matrix at COG w/ marginx[mm]

y[mm]

z[mm]

lxx[kg.m2]

lxy[kg.m2]

lxz[kg.m2]

lyy[kg.m2]

lyz[kg.m2]

lzz[kg.m2]

13.13 5.62 2.99 87.53 0.1687 0.0034 0.0030 0.57 -0.0087 0.0904

Table 4.24: S/C Mass, CoG and Inertia with component margins in Deployed config.

Z

Y

XZ

Y

X

Figure 4.10: S/C Stack Configuration (left) and S/C Deployed configuration (right).

43

Chapter 4. Subsystems: Analysis and Design

Magnetorquer

4xRW Pack

2 x PDU & ACU: PMAD

S/C OBC & UHF TRX

Battery Pack

Payload

UHF Antenna

StarTrackerSunSensor 1

SunSensor 2

SunSensor 3

Payload OBC &

S-Band TRX

S-Band Antenna

S/C Structure

Figure 4.11: S/C layout.

4.12 System Budgets and CubeSat DeployerTable 4.25 contains the detailed S/C mass budget. All COTS masses have been taken from their corre-sponding datasheets provided by their manufacturers (the mass indicated for the structure already includesthe secondary structure mass).

As can be seen, the total dry mass of the spacecraft without any margin is 11.5 kg. The payload is themain contributor to that value, making the overall mass of the S/C slightly above of a typical 6U CubeSat.On top of that value and according to the margin philosophy dictated by ESA and the product categoriesdescribed in the ECSS-E-ST-10-02C, 5% (Category A/B products, COTS TRL9), 10% (Category C, COTSwith minor modifications) or 20% (Category D, new design like the PL) margin is added as componentmargin. Besides, in this phase of the project a 20% system margin should be added. Hence, to total S/Cmass with all margins is 15.8kg.

Since the majority of CubeSat’s components are COTS, it could be expected that the real final totalmass will be between 11.5 and 14 Kg, which is the around the maximum mass that a wide number ofCOTS 6U CubeSat deployers, can afford. Nevertheless, if the final mass budget is above that values,some alternatives have been reviewed as a backup. There are certain deployers that allow certain type ofcustomisation like the ones offered by EXOLAUNCH company5).

Finally, Table 4.26 summarises the expected procumerement cost budget. Some manufactures publishthe cost procurement cost in their websites, but the majority of them only provides this informationupon request. The cost of those components has been estimated from homologous components of othermanufacturers. These cases are identified as "estimated" in the last column of the table.

5https://www.exolaunch.com/Products.html

44

4.12 System Budgets and CubeSat Deployer

Item QuantityUnitary Mass 

[kg]Total Mass 

[kg]Margin

Total mass with margin [kg]

Percentage of dry mass

COTS name

Structure 1,452 1,54 11,7%Structure 6U 1 1,1 1,100 5% 1,16 8,8% ISIS STR ‐ 6UShielding 1 0,352 0,352 10% 0,39 2,9%EPS 1,718 1,80 13,7%Solar array 6 0,2 1,200 5% 1,26 9,6% DEPSOLAR (MMA Design)Battery 1 0,27 0,270 5% 0,28 2,2% Gomspace ‐ NanoPower BP4 G OptionPMAD Motherboard 1 0,08 0,080 5% 0,08 0,6% P60 DOCKPDU 2 0,057 0,114 5% 0,12 0,9% Gomspace ‐ PDU ‐ 200ACU 1 0,054 0,054 5% 0,06 0,4% Gomspace ‐ ACU ‐ 200Communications 0,442 0,46 3,5%S‐Band Transceiver 1 0,06525 0,065 5% 0,07 0,5% Gomspace TR‐600S‐Band Antenna 1 0,11 0,110 5% 0,12 0,9% Gomspace ‐ Nanocom ANT6FS‐Band MotherBoard 1 0,07635 0,076 5% 0,08 0,6% Gomspace ‐ NanoDock DMC‐3UHF MotherBoard 1 0,07635 0,076 5% 0,08 0,6% Gomspace ‐ NanoDock DMC‐3UHF Transceiver 1 0,0245 0,025 5% 0,03 0,2% Gomspace NanoCom AX100 TransceiverUHF Antenna 1 0,09 0,090 5% 0,09 0,7% Gomspace ‐ Nanocom ANT2000ADCS 0,481 0,51 3,9%Reaction wheels Package 1 0,23 0,230 7% 0,25 1,9% 4x (RW210 HT)Star Tracker 1 0,04 0,035 5% 0,04 0,3% HT ‐ ST 200Sun Sensor 3 0,007 0,020 5% 0,02 0,2% Solar Mems ‐  Nano SSOC D60Magnetorquers 1 0,196 0,196 5% 0,21 1,6% ISIS ‐ MagnetorquerOBC 0,077 0,11 0,8%PL OBC 1 0,0768 0,077 5% 0,08 0,6% Gomspace ‐ Z7000SpaceCraft OBC 1 0,024 0,024 5% 0,03 0,2% Gomspace ‐ A3200Harness 0,200 0,24 1,8%Harness 1 0,2 0,200 20% 0,24 1,8%TCS 0,050 0,06 0,5%Thermal hardware 1 0,05 0,050 20% 0,06 0,5%Payload 7,000 8,40 64,0%PL 1 7 7,000 20% 8,40 64,0%Total Dry Mass 11,444 13,13 100,0%

20% 15,75Total Dry Mass w/  system margin

Table 4.25: System Mass Budget.

COTS name Quantity Unitary Price Total Price TRL Exact/Estimated PriceStructure 8.850 € ISIS STR - 6U 1 7.850 € 7.850 € 9 EstimatedShielding 1 1.000 € 1.000 € EstimatedEPS 34.000 € DEPSOLAR (MMA Design) 3 9.000 € 27.000 € 9 EstimatedGomspace - NanoPower BP4 G Option 1 3.000 € 3.000 € 9 EstimatedP60 DOCK 1 4.000 € 4.000 € 9 EstimatedGomspace - PDU - 200 2 - € Included in P60 DockGomspace - ACU - 200 1 - € Included in P60 DockCommunications 20.500 € Gomspace TR-600 1 8.500 € 8.500 € 9 EstimatedGomspace - Nanocom ANT6F 1 2.500 € 2.500 € 9 EstimatedGomspace - NanoDock DMC-3 2 500 € 1.000 € 9 EstimatedGomspace NanoCom AX100 Transceiver 1 3.500 € 3.500 € 9 EstimatedGomspace - Nanocom ANT2000 1 5.000 € 5.000 € 9 EstimatedADCS 61.700 € 4x (RW210 HT) 3 4.300 € 12.900 € 9 EstimatedHT - ST 200 1 30.000 € 30.000 € 9 EstimatedSolar Mems - Nano SSOC D60 3 3.600 € 10.800 € 9 ExactISIS - Magnetorquer 1 8.000 € 8.000 € 9 ExactOBC 5.800 € Gomspace - Z7000 1 5.800 € 5.800 € 9 EstimatedGomspace - A3200 1 2.900 € 2.900 € 9 EstimatedHarness 200 € Harness 1 200 € 200 € TCS 500 € Thermal hardware 1 500 € 500 €

134.450 €

Table 4.26: Procurement Cost Budget.

45

5. Risk Analysis and Mitigation

Table 5.1 shows the risk and mitigation analysis performed. Every identified risk is studied, assessingits impact to the mission. A severity number from 1 to 5 (being 1 the lowest severity) is multiplied by alikelihood number also from 1 to 5 (being 1 the lowest likelihood) to rank each risk. For those risks with ascore larger or equal to 15, a red colour appears. On the contrary, those risks exhibiting a 5 or lower scoreare highlighted in green colour.

For all risks, one or several mitigation actions are taken or proposed. The majority of these actionshave been already taken and included in the final design (see current status column). However, for abetter assessment of certain risks, further studies that are out of the scope of a PDR are required (ToBe Performed in the current status column). Nevertheless, the those required detailed analysis/test areproposed. This is the case of a EMC analysis, detailed TMM or TID analysis. In addition, the action’sdue date is included. Phase C and Phase D refers to the well-known ESA programme phases: DetailedDefinition (Phase C) and Production/Ground Qualification/testing (Phase D).

All risks marked in red are not accepted because their occurrence might impact significantly the accom-plishment of mission objectives or even S/C survivability. Important mitigation actions have been takento mitigate/eliminate them. In addition, the PL introducing biological material on Earth risk is alsonot accepted despite having a low score (determined by its low likelihood). At the end of the sciencephase, a command will be sent to kill the worms and associated viruses and bacteria. On the otherhand, an specific PL Control Plan will be elaborated in Phase C to set the procedures of manufacturing,manipulation, integration and testing of the PL. On top of that, exhaustive testing will be performed inPhase D at PL and S/C level to ensure that the payload will not be not damaged in ground handling andlaunch phases (random vibration test, sine test, shock test at PL and S/C level).

In addition to the risks stated, there are more external risks that shall be considered in future and moredetailed mission phases:• The CubeSat will be sent as a secondary payload in a launcher. The launcher selection depends on

external factors that are out of the scope of this PDR. The launcher performance and capabilities(like the launch load), will impact the design. Hence, when the main launcher and back-up launcherare selected, the CubeSat will be designed to withstand the envelope o both launchers’ loads.• Solar Activity: unexpected events related to the solar activity like CME. Since the mission duration

is short (3 months) in comparison with the period of solar activity (about 11 years), the launch datecan take into account this factor and adapt it to the minimum solar activity.• Depending on the procurement policy or programme foundation, some restrictions might apply for

certain non-european components (ECI policy1) or cost budget.• Risk of collision in the selected orbit: circular LEO are widely populated orbit. A space debris

analysis will be carried out un future phases to minimise the probability of collision.

1http://www.esa.int/Enabling_Support/Space_Engineering_Technology/European_Component_Initiative_ECI

Risk characterization Risk Mitigation StatusRisk Description

Title Description Impact

Severity

Likelyhood

Rank

Accep. Mitigation Action PlanCurrent Status

Due date

G/S not available in satellite overflight

A G/S considered on the coverage analysis is not available in an overflight of the CubeSat and PL data cannot be transmitted

Potential loss of PL data 5 3 15 NO

A1) Provide the CubeSat with a high-storage device to at least store 5 days of PL data

A1) Completed

PDR

S-Band module temporally inoperative

The S-Band module is not working temporally and communication cannot be achieved

Any mean for contacting the satellite

5 1 5 YES

A1) Utilize a reliable UHF module (transceiver+antennas) as a primary system for telemetry, tracking and command to recover S-Band module for operation

A1) Completed

PDR

S/C temporally attitude loss I

The S/C losses the Sun-pointing attitude and thus the power generation capacity

Lack of power generation

4 2 8 YES

A1) Equip a battery to at least support Safe Mode during 1hA2) Equip Sun Sensors to rapidly locate Sun position

A1) Completed A2) Completed

PDR

S/C temporally attitude loss II

The S/C losses the semi-Nadir pointing with -Z axes where the antennas are placed

Lack of comms

4 2 8 YES

A1) Use a set of 4 UHF omnidirectional antennas that allow transmiting and commanding the S/C regardeless of the attitude

A1) Completed

PDR

StarTracker big baffle required

A baffle is required for the StarTracker to operate in nominal conditions to reduce incoming light

The S/C config has to be changed

2 4 8 YES

A1) Leave some extra space under the StarTrackerA2) Positioning the StarTracker in the upper face which in nominal operation its away from light disturbances.

A1) Completed A2) Completed

PDR

EMC issues between the magnetorquer and the OBCs

The magnetorquer coils can affect the behaviour of the computers inside the satellite

Failure of OBC

5 3 15 NO

A1) Positioning the OBCs and the magnetorquer away from each otherA2) Perform a EMC study to better evaluate the issue.

A1) CompletedA2) To Be performed

A1) PDRA2) Phase C

RW not working

One reaction wheel fails jeopardizing S/C attitude control

Loss of 3 axis attitude control 5 1 5 YES

A1) An extra reaction wheel is equipped on top of the 3 minimum required one for 3 axis controlA2) A backup solution of using magnetorquers for attitude control

A1) Completed A2) Completed

PDR

TMM assumptions

The TMM has been created with several assumptions due to the lack of COTS thermal properties and does not match real properties

Temps out of the Design ranges 5 4 20 NO

A1) Qual (5°C), Accept(5°C), and uncertainty(10°C) margins have been included in the analysis wrt the COTS qual tempsA2) Detailed Thermal Analysis and Tests

A1) CompletedA2) To Be performed

A1) PDRA2) Phases C and D

PL introducting worms/virus/bacteria on Earth

The biological PL or associated viruses or bacterias introduced on Earth in any part of the mission lifecycle

Earth contamination

5 1 5 NO

A1) Command the PL to kill worms, virus and bacteria after science phaseA2) Set specific control plan for PL manufacturing, manipulation and testA3) Testing at PL and S/C level

A1) CompletedA2) To be performedA3) To be performed

A1) PDRA2) Phase CA3) Phase D

PL size larger than defined one

The final design of the PL occupies a larger volume than the 3U defined one

S/C volume budget 3 3 9 YES

A1) There is 0.7U volume margin to cover a potential larger payload

A1) Completed

PDR

Larger torque disturbance

The estimated external torque perturbations are lower than the real ones and other torque sources appear PL pumping, EM torques induced in electronic devices, etc.

Attitude Control Sizing

2 3 6 YES

A1) Select COTS that can counteract at least twice the external torque estimationA2) Flexible S/C design to allow for changes

A1) Completed A2) Completed

PDR

Thermal shock torques

In orbit thermal shocks that affect flexible appendages genrating extra-torques

Attitude Control

3 1 3 YES

A1) This problem is intrinsicly mitigated with the eclipse-free orbit selection

A1) Completed

PDR

Battery Malfunction

Sudden failure of the battery NOM+COM operational mode 4 1 4 YES

A1) The S-Band antenna would be operated in a lower regime reducing the gainA2) Rely on TRL9 COTS with wide flight heritage

A1) Completed

PDR

Radiation shielding not enough

The TID that COTS can receive is lower than the one expected

Electronics damage, SEE 5 3

15 NOA1) Perform a detailed TID analysis

A1) To Be Performed

Phase C

Table 5.1: Risk and Mitigation analysis.

47

6. Summary and Conclusions

This Preliminary Design Report has presented the achieved CubeSat design solution to perform theambitious Roscoff Worms Mission. To reach this target, first, a clear definition of the scope of this workand management strategy were set. A key point was the derivation of lower level requirements fromthe mission statement and the establishment of the design philosophy: find the most robust, maturedand reliable solution that complies with the payload needs and requirements. To do so, the proposedsolution only relies on COTS equipment, avoids the use of a propulsion system (which has the lessmatured technologies for CubeSats and would imply a significant risk for the payload) and minimisesorbit geometry and system complexity.

An iterative concurrent design process has been carried out in the framework of both the top level require-ments and derived requirements. The Star Worms UC3M team has utilised concurrent design tools like theIDM-CIC and the System Data Summary to converge the design process. For all S/C subsystems a COTSmarket research has been done to perform trade-offs whose outcome contributes to the optimisation ofthe design. A definition of the mission and system architecture was performed to ease the comprehensionof the interaction among the different segments and/or systems.

The result of the mission analysis is a 500 km-altitude SSO. Thanks to this orbit, eclipse periods areeliminated (with a 5 months per year launch window and 3 months mission duration). Besides, it allowsa more stable thermal environment which has ease the thermal design. Furthermore, the SSO enablessteady power generation (absence of eclipses) which has reduce the EPS design complexity. Moreover,the altitude is selected to comply with the end-of-life regulation since the CubeSat naturally de-orbitsand burns in the atmosphere before 25 years. Last but not least, the the mission analysis included a G/Scoverage analysis which has demonstrated that the CubeSat is able to transmit daily all the generated PLdata to the ESA CORE Network.

The satellite will communicate with the ground stations through two transmitter-antenna pairs in orderto increase the robustness of the communications, one S-band and one UHF antenna. The first one isused for transmitting payload data, due to its high data rate capacity; while the latter is used to transmitspacecraft telemetry and receive telecommands, due to having an omnidirectional antenna, allowing com-munication even under severe pointing misalignment. The specific components for each pair were selectedby performing trade-off studies, and a RF link budget was performed for every link in order to ensurethe communication subsystem was capable of meeting the requirements. To manage the transceivers, thesatellite will have two on-board computers, a main one for housekeeping, telemetryand commands; and adedicated payload computer for image processing and storage.

Power generation is ensured using a partially-deployable solar array (42 cells for 44.47 W EOL powergeneration capacity) and continuous sun-pointing. The size of the solar array was determined by thepower consumption in nominal mode (payload operation but no data transmission to Earth) since thecommunication windows along a day represent only a small fraction of time (≈ 1/24) and the powerrequired by communication system was about 1/3 of the total required power. Hence, to avoid over-sizingthe solar arrays, the extra power required on the nominal+communications operational mode (payloadoperation and data transmission at the same time) are provided by the battery. Finally, a PMAD is incharge of receiving, managing and distributing the power among the different units.

In terms of attitude determination and attitude control the system requirements have been fully accom-plished. The solution to compensate the environmental disturbances consist on four reaction wheelsset in a pyramidal structure. The additional wheel provides robustness to the attitude control of the S/C.Moreover, a magnetic torque is located in order to allow the reaction wheels to desaturate and it provides

6.1 Future Works

redundancy in terms of CubeSat detumbling and attitude control since the mission operates at LEO. Theattitude determination is provided by a star-tracker and three sun sensors giving, as a result, an accurateand redundant solution.

Regarding the TCS, in order to maintain the units in their required temperature range, an external insula-tion has been installed. On top of that, the PL has shown to need further insulation. For this reason, ithas been designed to be covered with MLI. For the sake of ensuring stable temperatures and reduce TCScomplexity, specially in a possible cold scenario, only passive systems and heaters have been proposed.The difficulty of the problem has led to some complex decisions and several iterations on the design that,at the end, have kept components’ temperature within their design range.

All the previous systems are encompassed in a 6U CubeSat structure covered by aluminium panels forradiation shielding. The overall mass of the CubeSat with margins is 15.75 kg. The CubeSat layout wasdesigned to allow steady power generation at the same time the CubeSat communicates with Earth. Thesummary of the mission and the COTS selected is included in Table 6.1. Finally, during the whole designprocess risks were identified and mitigation actions were taken to minimise or even eliminate theirimpact.

6.1 Future Works

This PDR has demonstrated that the achieved solution is a robust, mature and reliable design. Nevertheless,more detailed analysis can be performed with this work as a basis to further develop the CubeSat design.A list of future works are listed below (still for design phase):• In general, contact the manufacturers of the selected COTS components to request more detailed

information that allows a more detailed design.• EMC analysis to ensure that the magnetorquer will not affect the rest of S/C electronic components.• A detailed solar array power generation capacity study. Solar array’s temperature affects cells

performance which also affects the total power available. This studies should be performed involvingthe SA manufacturer.• Radio-Frequency Analysis: a RF analysis can be performed to study the interaction of S/C compo-

nents with antennae performance.• Communications detailed scheduling: forecast the data generation and transmission and set a

detailed PL data transmission plan. With this, the use of G/S can be further optimised, thus reducingoperating costs.• Create a TMM with ESATAN software to run thermal cases and improve the temperature evolution

estimation. Apart from the studied nominal case and cold case, other cases like the deploymentphase case shall be added• Since the PL is the most relevant constraint of the TCS system, further study needs to be done to

find the most efficient design. A finite elements study of the temperature distribution could provideunderstanding of the temperature requisites. Probably, it is possible not to maintain the whole PLbetween 11 and 15 oC, but only the volume containing the Roscoff worm colony. This fact wouldrelax the constrains, and allow a more reliable and less complex design.• Environmental effects analysis: TID analysis and SEE analysis for electronic components and

straight-light analysis for optical components. Outgassing analysis.• Space debris analysis to evaluate and minimise the probability of in-orbit collision.• Launcher analysis: evaluate the loads introduced by the selected launcher in terms of random loads,

sine loads and shock loads.• Star the development of the Fault Detection, Isolation and Recovery (FDIR): this work involves

forecasting potential failures and developing decision trees to recover nominal operation if theyoccur.

49

Chapter 6. Summary and Conclusions

• Propagate and analyse the environmental disturbances during the whole time of operation. Thiswill allow to reinforce the principal axis of inertia with the actuators.• Adjust the size of the sensors of ADCS with vibration tests to achieve accurate solution.• In terms of mechanical design, structural analysis of appendages and deployable parts. FEM

analysis will be mandatory.

Orbit ‐ SSO h=500km circular

Lifetime ‐ Nominal operational lifetime: 3 months

Launch window ‐ From 8th February to 1st August

Mass ‐ Dry mass: 15.75 kg (with margins)

Dimenssions‐ Stowed: 320x210x110 mm

‐ Deployed (solar panels): 320x600x110 mm

Payload‐ Carries living Roscoff Worms 

‐ L‐shaped 3U box of 7kg

ADCS

Actuators:

     ‐ 4 Hyperion RW210 6.0 mNms reaction wheels

     ‐ 1 ISIS magnetorquer board

Sensors:

      ‐ 3 Solarmems nanoSSOC‐D60 sun sensors with 60° FOV

      ‐ 1 Hyperion ST200 star tracker with 30 arcseconds accuracy

Communications

‐ 1 GOMspace NanoCom TR‐600 S‐Band transceiver

‐ 1 GOMspace NanoCom ANT2000 S‐Band antenna

‐ 1 GOMspace NanoCom AX100 UHF transceiver

‐ 1 GOMspace NanoCom ANT‐6F UHF antenna

OBC‐ S/C OBC: GOMspace Nanomind A3200

‐ Payload OBC: GOMspace Nanomind Z7000 with 32GB capacity

EPS

‐ Deployable solar arrays: MMA Design + 32% eff Azurspace cells

‐ 1 GOMspace NanoPower BP4 battery with 38,5 Wh capacity

‐ 1 GOMspace NanoPower P60 PMAD motherboard with

  2xPDU and 1xACU

TCS‐ Heaters and condution connections

‐ External MLI and PL MLI

Structure ‐ ISIS 6‐Unit CubeSat structure

Mission Description

Overall system characteristics

Payload and Subsystems

Table 6.1: Mission and S/C characteristics.

50

A. Mission Requirements and Compliance

This appendix contains the top level requirements taken from the mission statement as well as the derivedlower level requirements. In addition, the parent requirement, the verification method and the location ofthe compliance justification for those requirements verified by analysis in this included.

1. Top Level Requirements: there are 10 top level requirements listed in the mission statementdenoted with the following identification structure: RWX (where X is the requirement number).

2. System/Orbit Level Requirements: they are derived from the Top Level Requirements. Hence,in their parent requirement column always appear a "RWX". Their ID is composed of thee terms:• NS: NanoStar.• SYS/ORB: SYS for System requirements and ORB for Orbit requirements.• YY: two digits number.

3. Subsystem Level Requirements: they are derived from the System/Orbit Level Requirements.Hence, in their parent requirement column always appear either NS-SYS-XX or NS-ORB-XX. TheirID is composed of thee terms:• NS: NanoStar.• ZZZ: Subsystem identifier. ADC (Attitude Determination and Control Subsystem), CDH

(Command and Data Handling), COM (Communications Subsystem), EPS (Electrical PowerSystem), STR (Structure subsystem) and TCS (Thermal Control Subsystem).• YY: two digits number.

Table A.1: Requirements list, Parent requirement and Verification.

ID Requirement Parent Verif.Method

Sec.

RW1 The system shall carry, activate and operate safelythe main scientific payload described below, duringthe maximum time in orbit, thus maximizing the totalamount of scientific data received on Earth

N/A Review ofDesing

RW2 The scientific data obtained by the main science pay-load shall be transmitted to Earth within the missiontime-frame

N/A Review ofDesing

RW3 The satellite shall be capable of performing the mis-sion objectives, considering the space environmentconstraints

N/A Review ofDesing

RW4 The satellite shall guarantee the correct attitude ofthe main science payload (depending on the chosenpayload design) or the communication system duringthe orbits whenever needed

N/A Review ofDesing

RW5 The satellite volume shall not exceed 8U N/A Review ofDesing

RW6 The mission duration from launch to end-of-life shallnot be lower than 2 weeks and ideally up to 3 months.

N/A Review ofDesing

RW7 The maximum mission duration from launch to end-of-life shall comply with space regulation

N/A Review ofDesing

RW8 Ground segment(s) shall rely only on the trackingstations of the ESA network

N/A Review ofDesing

Chapter A. Mission Requirements and Compliance

ID Requirement Parent Verif.Method

Sec.

RW9 During all its lifecycle, the nanosatellite shall notintroduce any of its embedded photosymbiotic flatworms or its associated viruses or bacteria populationinto any area on Earth

N/A Review ofDesing

RW10 The conception of the nanosatellite shall ensure thatall its lifecycle tooling and environment is ROHS (Re-striction Of Hazardous Substances in electrical andelectronic equipments), REACH (Registration, Eval-uation, Authorization and Restriction of Chemicals)and CSR (Corporate Social Responsibility) compli-ant

N/A Review ofDesing

NS-ORB-01 The orbit shall be selected to allow the transmissionto Earth of the 550MB of scientific data generatedper day by the PL within the mission time-frame

RW2 Analysis 4.1

NS-ORB-02 The Routine Operations Phase of the mission shalllast a minimum of 2 weeks

RW6 Analysis 4.1

NS-ORB-03 The satellite shall de-orbit before 25 years after theend of S/C operations

RW7 Analysis 4.1

NS-ORB-04 The satellite shall burn in the atmosphere after itsoperational life

RW9 Review ofDesing

NS-SYS-01 The system shall be capable of carrying a 7kg-3U-L-shaped payload

RW1 Review ofDesing

NS-SYS-02 The system shall be capable of providing 9.5W to thepayload

RW1 Review ofDesing

NS-SYS-03 The system shall be capable of transmitting to Earththe 550MB of scientific data generated per day bythe PL within the mission time-frame

RW2 Analysis 4.1

NS-SYS-04 The satellite shall proctect S/C components againstthe impact of electrons and ions present in the se-lected orbit

RW3 Analysis 4.9.2

NS-SYS-05 The satellite shall have the capability of keeping S/Ccomponents within their Design Temperature range

RW3 Analysis/Test

4.10.2

NS-SYS-06 The satellite shall have attitude determination andcontrol capabilities

RW4 Review ofDesing

NS-SYS-07 The satellite shall never exceed a maximum rotationspeed of 3RPM for more than 10 s

RW4 Analysis 4.8.2

NS-SYS-08 The satellite shall accomodate all componets (includ-ing the PL) in a structure of less than 8U

RW5 Review ofDesing

NS-SYS-09 The satellite shall transmit all the data only to thestations of the ESA CORE Network

RW8 Review ofDesing

NS-SYS-10 The satellite shall be commanded only from the ESACORE Network

RW8 Review ofDesing

NS-SYS-11 The satellite shall has the capability of commandingthe payload

RW9 Review ofDesing

NS-SYS-12 The system shall not have a Propulsion Subsystemneither for primary propulstion nor for attitude con-trol

RW6 Review ofDesing

52

ID Requirement Parent Verif.Method

Sec.

NS-SYS-12 The conception of the nanosatellite shall ensure thatall its lifecycle tooling and environment is ROHS(Restriction Of Hazardous Substances in electricaland electronic equipments)

RW10 Review ofDesing

NS-SYS-13 The conception of the nanosatellite shall ensure thatall its lifecycle tooling and environment is REACH(Registration, Evaluation, Authorization and Restric-tion of Chemicals)

RW10 Review ofDesing

NS-SYS-14 The conception of the nanosatellite shall ensure thatall its lifecycle tooling and environment is CSR (Cor-porate Social Responsibility) compliant

RW10 Review ofDesing

NS-SYS-15 Declared materials, mechanical parts and processeslists shall be produced according to the Declaredmaterial list (DML) DRD specified in ECSS-Q-ST-70

RW9 Review ofDesing

NS-SYS-16 The satellite shall survive to the launch phase withoutdamaging the payload

RW9 Review ofDesing

NS-ADC-01 The AOCS shall be capable of determining its atti-tude with a pointing accuracy of 0.5 deg during theRoutine Operations Phase of the mission

NS-SYS-06 Analysis 4.8.2

NS-ADC-02 The AOCS shall be capable of determining S/C at-titude with a pointing accuracy of 5 deg in the SafePower mode

NS-SYS-06 Analysis 4.8.2

NS-ADC-03 The RW shall be capable of providing continouslya slew torque of 1 · 10−4 mNm during 24h beforedesaturation

NS-SYS-06 Analysis 4.8.2

NS-ADC-04 The RWs shall be capable of counteracting the exter-nal torques continously during 24h before desatura-tion

NS-SYS-06 Analysis 4.8.2

NS-ADC-05 The AOCS shall be capable of keeping the rotationalspeed of the satellite bellow 3RPM from launcherseparation to the end of the Routine Operations Phaseof the mission

NS-SYS-07 Analysis 4.8.2

NS-CDH-01 The C&DH shall be capable of storing the PL datacoresponding to 10 days of operation

NS-SYS-03 Review ofDesing

NS-CDH-02 The C&DH System shall be able to process com-mands sent from the ESA CORE Network

NS-SYS-10 Review ofDesing

NS-CDH-03 The C&DH System shall be able to send a commandto the PL to eliminate all virus/bacteria after the Rou-tine Operations Phase of the mission

NS-SYS-11 Analysis/Test

4.6

NS-COM-01 The Communication System shall be capable of trans-mitting to the Earth the 550MB of scientific datagenerated in a day during that day

NS-SYS-03 Analysis 4.1

NS-COM-02 The Commmunications System shall be compatiblewith the antennas of the ESA CORE Network

NS-SYS-09 Review ofDesing

53

Chapter A. Mission Requirements and Compliance

ID Requirement Parent Verif.Method

Sec.

NS-COM-03 The Commmunications System achieve the PL datatransmission with a margin of at least 2dB with re-spect to the required Energy per Bit Noise Ratio

NS-SYS-03 Analysis 4.5.3

NS-COM-04 The Commmunications System achieve the PL datatransmision with a maximum BER of 10E-6

NS-SYS-03 Analysis 4.5.3

NS-EPS-01 The EPS shall provide 0.1W to the PL-Camera whenin SUSPEND mode from launcher separation to theend of the Routine Operations Phase of the mission

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-02 The EPS shall provide 5W to the PL-Camera whenswitched ON during the Routine Operations Phaseof the misssion

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-03 The EPS shall provide continuously 1W to the probesof the PL from launcher separation to the end of theRoutine Operations Phase of the mission

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-04 The EPS shall provide continuously 0.5W to thepumping system of the PL from launcher separationto the end of the Routine Operations Phase of themission

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-05 The EPS shall be capable of providing continuously2W to thePL heater from launcher separation to theend of the Routine Operations Phase of the mission

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-06 The EPS shall be capable of providing continuously1 W to the PL LED from launcher separation to theend of the Routine Operations Phase of the mission

NS-SYS-02 Review ofDesign/Analysis

4.7.3

NS-EPS-07 The EPS shall have the capability of providing con-tinuosly 10 W to the TCS

NS-SYS-05 Review ofDesign/Analysis

4.7.3

NS-STR-01 The Structure shall include an aluminum layer of1mm of equivalent thickness

NS-SYS-04 Review ofDesing

NS-STR-02 The structure shall withstand launcher frequenciesup to 300Hz

NS-SYS-16 Analysis 4.9.1

NS-TCS-01 The Thermal Control strategy shall rely only on coat-ings, heaters and radiators

NS-SYS-05 Review ofDesing

NS-TCS-02 The mission phases shall be represented by a coher-ent set of thermal design cases covering the extremerange of conditions experienced in the mission life-time

NS-SYS-05 Review ofDesing

NS-TCS-03 The TCS shall consider a PL LED illumination pro-file of 2.5h ramp-up, 1h flatbed and 1.5h ramp-down

NS-SYS-02 Review ofDesing

54

TOP LEVEL SYSTEM/ORBIT LEVEL SUBSYSTEM LEVEL

NS‐SYS‐01

NS‐EPS‐01

NS‐EPS‐02

NS‐EPS‐03

NS‐EPS‐04

NS‐EPS‐05

NS‐EPS‐06

NS‐TCS‐03

NS‐COM‐01

NS‐CDH‐01

NS‐COM‐03

NS‐COM‐04

NS‐ORB‐01

NS‐SYS‐04 NS‐STR‐01

NS‐EPS‐07

NS‐TCS‐01

NS‐TCS‐02

NS‐ADC‐01

NS‐ADC‐02

NS‐ADC‐03

NS‐ADC‐04

NS‐SYS‐07 NS‐ADC‐05

RW5 NS‐SYS‐08

NS‐ORB‐02

NS‐SYS‐12

RW7 NS‐ORB‐03

NS‐SYS‐09 NS‐COM‐02

NS‐SYS‐10 NS‐CDH‐02

NS‐SYS‐11 NS‐CDH‐03

NS‐SYS‐15

NS‐SYS‐16 NS‐STR‐02

NS‐ORB‐04

NS‐SYS‐12

NS‐SYS‐13

NS‐SYS‐14

NS‐SYS‐05

RW4NS‐SYS‐06

RW1NS‐SYS‐02

RW2NS‐SYS‐03

RW10

RW6

RW8

RW9

RW3

Table A.2: Requirements Flowdown Table

55

B. N2chart

The N2 chart is a diagram in the shape of a matrix, representing functional or physical interfaces betweensystem elements. It is used to systematically identify, define, tabulate, design, and analyze functionaland physical interfaces. It applies to system interfaces and hardware and/or software interfaces1. Thefunctioning of this tool is shown in Figure B.1

Figure B.1: N2 chart functioning diagram.

1https://en.wikipedia.org/wiki/N2_chart

Orb

itCo

mm

s op

port

uniti

es:

cove

rage

Orb

it po

sitio

n,

eclip

ses

S/C

posi

tion;

De-

orbi

tS/

C at

titud

e:

Sun

poin

ting

S/C

attit

ude:

Su

n po

intin

g

S/C

posi

tion

in

mis

sion

pha

ses;

D

e-or

bit

Spec

ific

Spac

e En

viro

nmen

t

S/C

att

itude

; Ec

lipse

s; S

olar

ra

d., A

lbed

o &

pl

anet

ary

IR

PayL

oad

Dat

a ac

quis

ition

Pow

er D

eman

dSy

stem

sta

tus

Syst

em s

ize,

sh

ape

& m

ass

Ope

ratin

g Te

mp

Payl

oad

Proc

esso

rSt

ored

PL

data

tr

ansm

issi

onPo

wer

Dem

and

Syst

em s

tatu

sSy

stem

siz

e,

shap

e &

mas

sO

pera

ting

Tem

p

Com

ms

Pow

er D

eman

dSy

stem

sta

tus

Poin

ting

accu

racy

re

quire

d

Tele

met

ry a

nd

PL d

ata

tran

smis

sion

G/S

dat

a pr

oces

sing

Syst

em s

ize,

sh

ape

& m

ass

Ope

ratin

g Te

mp

Pow

er

gene

ratio

n st

rate

gy

Ante

nna

feed

ing

pow

erEP

SSy

stem

sta

tus

Poin

ting

accu

racy

re

quire

d

Syst

em s

ize,

sh

ape

& m

ass

Ope

ratin

g Te

mp

Tele

met

ry d

ata

tran

smis

sion

Tele

met

ry

In-o

rbit

attit

ude

cont

rol

Pow

er D

eman

dSy

stem

sta

tus

ADCS

Actu

ator

s an

d se

nsor

s da

taSy

stem

siz

e,

shap

e &

mas

sO

pera

ting

Tem

p

Cove

rage

tim

eTT

&C

tran

smis

sion

Gro

und

Segm

ent

Traj

ecto

ry

trac

king

Nav

igat

ion

Pow

er D

eman

dSy

stem

sta

tus

GN

C

Trac

king

Dat

a;

Com

man

ds to

un

its; P

L co

ntro

l

Syst

em s

ize,

sh

ape

& m

ass

Ope

ratin

g Te

mp

Com

man

ds to

PL

Hou

seke

epin

g D

ata

Tran

ssm

isio

n

Pow

er D

eman

d;

oper

atio

nal

mod

es s

witc

hSy

stem

sta

tus

Attit

ude

cont

rol

law

Traj

ecto

ry

prop

agat

ion

S/C

OBC

Syst

em s

ize,

sh

ape

& m

ass

Ope

ratin

g Te

mp

Bus

pow

er

dem

and

Syst

em s

tatu

sM

ass

& In

ertia

Pr

oper

ties

Stru

ctur

e &

Co

nfig

urat

ion

Tem

p ra

nge.

Th

erm

al

stre

sses

Ther

mal

con

trol

Ther

mal

con

trol

Ther

mal

con

trol

Pow

er D

eman

d;

Ther

mal

Co

ntro

lSy

stem

sta

tus

Ther

mal

con

trol

Ther

mal

con

trol

Ther

mal

con

trol

Syst

em s

ize,

sh

ape

& m

ass

TCS

Syst

ems

inte

ract

ion

: N2

Ch

art

Figure B.2: N2 chart.

57

C. COTS Datasheet Links

Unit Link to DatasheetReaction Wheel https://hyperiontechnologies.nl/wp-content/uploads/2019/11/HT_

RW210.pdfStar Tracker https://hyperiontechnologies.nl/wp-content/uploads/2015/07/HT_

ST200_v2.1-flyer.pdfSun Sensors https://www.cubesatshop.com/wp-content/uploads/2016/06/

nanoSSOC-D60-Technical-Specifications.pdfBattery https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanopower-bp4-27.pdfOBC https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanomind-a3200_1006901.pdfPL-ADCS OBC https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanomind-z7000-15.pdfPMAD https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanopower-p60-dock-29.pdfStructure https://www.isispace.nl/wp-content/uploads/2019/08/

ISIS-CubeSat-Structures-Brochure-v2R-web_compressed.pdfMagentorquers https://www.isispace.nl/wp-content/uploads/2016/02/

iMTQ-Brochure-v1.pdfS-Band Transceiver https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanocom-tr600-16.pdfS-Band Antenna https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanocom-ant2000.pdfUHF Transceiver https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanocom-ax100.pdfUHF Antenna https://gomspace.com/UserFiles/Subsystems/datasheet/

gs-ds-nanocom-ant6f-uhf-21.pdf

Table C.1: COTS Datasheet links.

D. Project Management Charts

,

RW

Cu

be

Sat

Pro

ject

Pro

ject

Init

iati

on

Pro

ject

Def

init

ion

Co

nce

ptu

al D

esig

n:

Co

ncu

rren

t En

gin

eeri

ng

Det

aile

d D

esig

n:

Co

ncu

rren

t En

gin

eeri

ng

Pro

ject

Clo

sure

Team

org

aniz

atio

n:

-Se

tm

emb

ers

skill

s-

Team

role

sd

efin

itio

n

Pro

ject

Pla

nn

ing:

-M

issi

on

co

nce

pt

dis

cuss

ion

-M

issi

on

sta

tem

ent

-W

ork

Bre

akd

ow

n-

Wo

rk F

low

Init

ial

dat

a ge

ne

rati

on

: -

Lite

ratu

re r

evie

w-

Pre

limin

ary

syst

ems

sele

ctio

n-

Stat

e-o

f-th

e-ar

t o

verv

iew

-D

ata

colle

ctio

n

Re

qu

ire

me

nts

de

fin

itio

n:

-To

ple

velR

Qs

-M

issi

on

RQ

s-

Syst

ems

def

init

ion

-Sy

stem

leve

lRQ

s-

Sub

syst

emle

velR

Qs

Fun

ctio

nal

anal

ysis

:-

Mis

sio

no

verv

iew

-Fu

nct

ion

alfl

ow

-O

per

atio

ns

def

init

ion

Syst

em

s in

tegr

atio

n:

-N

2ch

art

Mis

sio

n c

on

cep

tual

an

alys

is:

-O

rbit

sel

ecti

on

-O

rbit

des

ign

-G

/S C

ove

rage

-Ec

lipse

ph

ases

-Sp

ace

Envi

ron

men

t-

Co

mm

un

icat

ion

s &

O

per

atio

ns

des

ign

-O

per

atio

ns

trad

e-o

ff-

Ch

eck

for

req

uir

emen

ts

com

plia

nce

Sub

syst

em

sco

nce

ptu

ald

esi

gn:

-Su

bsy

sco

nce

pts

-Su

bsy

sd

esig

n o

pti

on

s-

Trad

e-o

ff c

rite

ria

def

init

ion

-Tr

ade-

off

an

alys

is a

nd

mat

rix

-El

imin

atio

n o

f su

bo

pti

mal

o

pti

on

s-

Esta

blis

hm

ent

of

feas

ibili

ty-

Ch

eck

for

req

uir

emen

ts

com

plia

nce

Mis

sio

n d

eta

iled

an

alys

is:

-C

om

ple

te T

raje

cto

ry

pro

pag

atio

n-

An

alys

is o

f an

om

alie

s &

su

bo

pti

mal

op

erat

ing

con

dit

ion

s-

Full

eclip

se a

nd

G/S

co

vera

gean

alys

is

Sub

syst

em

sd

eta

iled

de

sign

:-

Det

aile

d c

on

curr

ent

sub

syst

ems

des

ign

an

d s

izin

g.-

CO

TS F

inal

Tra

de-

off

an

d

sele

ctio

n

Syst

em

sV

eri

fica

tio

n:

-C

hec

k fo

r re

qu

irem

ents

co

mp

lian

ce

Fin

al d

eliv

era

ble

s:-

Pre

limin

ary

Des

ign

Rep

ort

-P

roje

ct P

rese

nta

tio

n-

IDM

-CIC

file

De

sign

re

vie

w:

-C

hec

k d

etai

led

des

ign

an

d

veri

fica

tio

n d

ata

-D

esig

n R

evie

w m

eeti

ng.

-En

sure

sys

tem

s in

tegr

atio

n.

-A

nal

yze

and

iter

ate

des

ign

Fin

ald

esi

gnre

vie

w:

-C

he

ckd

etai

led

des

ign

san

dco

rrec

tio

ns.

-Fi

xm

ino

rch

ange

s-

Fin

aliz

eal

ld

esig

ns

and

clo

seo

ut

the

mis

sio

n.

Mis

sio

nan

dSy

ste

mLe

vel:

-Es

tab

lish

men

to

fFe

asib

ility

-Sy

stem

sA

rch

itec

ture

-R

isk

An

alys

is

Figure D.1: Project Work Break Down Structure.

Chapter D. Project Management Charts

Co

nce

ptu

al

des

ign

re

view

. C

lear

ance

?

Team

o

rgan

isat

ion

Mis

sio

n

con

cep

ts

dis

cuss

ion

Wo

rk

pla

nn

ing

&

bre

akd

ow

n

Init

ial d

ata

gen

erat

ion

an

d s

tate

-of-

the-

art

revi

ew

Mis

sio

n

app

licat

ion

s st

ud

y

Pre

limin

ary

syst

ems

ove

rvie

w

Pro

ject

init

iati

on

Mis

sio

n

def

init

ion

Mis

sio

n

Ove

rvie

w

Syst

em f

un

ctio

ns

anal

ysis

Fun

ctio

nal

flo

w

Syst

ems

inte

grat

ion

Pro

ject

Def

init

ion

Req

uir

emen

ts

def

init

ion

Mis

sio

n a

nal

ysis

Op

ers.

d

esig

n

Op

erat

ion

al

trad

e-o

ff

Op

tim

al

orb

it

des

ign

Req

s.

chec

k

Feas

ibili

ty

chec

kTr

ade-

off

p

roce

ss

Des

ign

sele

ctio

ns

Sub

syst

em

s d

esig

n

Sub

sys

con

cep

ts &

o

pti

on

s

Fin

al

sub

syst

ems

sele

ctio

n,

arch

itec

ture

, an

d

inte

grat

ion

Co

nce

ptu

al D

esi

gn:

con

curr

en

t e

ngi

ne

eri

ng

Mis

sio

n a

nal

ysis

Det

aile

d o

rbit

d

esig

n

An

alys

is o

f an

om

alie

s an

d

off

-no

min

al

con

dit

ion

s

Sub

syst

ems

des

ign

Det

aile

d

sub

syst

ems

des

ign

, si

zin

g, p

ow

& T

C

man

agem

ent

Det

aile

d s

yste

ms

inte

grat

ion

Syst

ems

Ver

ific

atio

n

Det

aile

d

des

ign

re

view

. C

lear

ance

?

No

Yes

Yes

No

Fin

al d

esig

n

revi

ew.

& p

roje

ct

clo

se-o

ut

Pro

ject

D

eliv

erab

les

Det

aile

d D

esi

gn:

con

curr

en

t e

ngi

ne

eri

ng

Pro

ject

Clo

sure

Figure D.2: Project Work Flow.

60

E. Coverage and Link Budget Charts

Figure E.1: Contacts with ESA CORE Network in 24h.

Figure E.2: Signal attenuation due to the atmosphere.

Chapter E. Coverage and Link Budget Charts

Figure E.3: Signal attenuation due rain.

Figure E.4: Energy per bit to noise ratio required for the desired BER.

62

F. Thermal Control Subsystem Charts

F.1 Model representation

Figure F.1: Representation of the radiation connections between nodes.

Figure F.2: Representation of the conduction connections between nodes

Chapter F. Thermal Control Subsystem Charts

F.2 Hot case: temperature evolution of all nodes along time

0 500 1000 1500

Time [min]

240

250

260

270

280

290

300

310

320

330

340

Tem

pera

ture

[K

]

Interior panels temperature

Front panel

Back panel

Right panel

Left panel

Top panel

Bottom panel

Figure F.3: Interior panels temperature.

0 500 1000 1500

Time [min]

150

200

250

300

350

400

Tem

pera

ture

[K

]

MLI temperature

MLI front panel

MLI back panel

MLI right panel

MLI left panel

MLI top panel

MLI bottom panel

Figure F.4: MLI temperature.

0 500 1000 1500

Time [min]

220

240

260

280

300

320

340

360

380

400

Tem

pera

ture

[K

]

Solar panels temperature

Central solar panel

Right solar panel

Left solar panel

Figure F.5: Solar panels temperature.

64

F.2 Hot case: temperature evolution of all nodes along time

0 500 1000 1500

Time [min]

240

260

280

300

320

340

360

Tem

pera

ture

[K

]

Antennas temperature

S-Band antenna

UHF Antenna

Figure F.6: Antennas temperature.

0 500 1000 1500

Time [min]

240

260

280

300

320

340

360

Tem

pera

ture

[K

]

Electronic devices temeprature

S-Band Transceiver

UHF Transceiver

SpaceCraft On-board computer

Payload On-board computer

Power management and distribution

Figure F.7: Electronic devices temperature.

0 500 1000 1500

Time [min]

260

265

270

275

280

285

290

295

300

305

310

Tem

pera

ture

[K

]

ADCS devices temperature

Magnetorquer

Reaction Wheels

Star Tracker

Figure F.8: Electronic controllers temperature.

65

Chapter F. Thermal Control Subsystem Charts

0 500 1000 1500

Time [min]

270

275

280

285

290

295

300

305

310

315

320

Tem

pera

ture

[K

]

Battery temperature

Battery

Figure F.9: Battery temperature.

0 500 1000 1500

Time [min]

284

284.5

285

285.5

286

286.5

287

287.5

288

288.5

Tem

pera

ture

[K

]

Payload temperature

Payload

Figure F.10: Payload temperature.

F.3 Cold case: temperature evolution of all nodes along time

0 500 1000 1500

Time [min]

220

240

260

280

300

320

340

360

380

400

Tem

pera

ture

[K

]

Solar panels temperature

Central solar panel

Right solar panel

Left solar panel

Figure F.11: Solar panels temperature.

66

F.3 Cold case: temperature evolution of all nodes along time

0 500 1000 1500

Time [min]

240

260

280

300

320

340

360

Tem

pera

ture

[K

]

Antennas temperature

S-Band antenna

UHF Antenna

Figure F.12: Antennas temperature.

0 500 1000 1500

Time [min]

240

260

280

300

320

340

360

Tem

pera

ture

[K

]

Electronic devices temeprature

S-Band Transceiver

UHF Transceiver

SpaceCraft On-board computer

Payload On-board computer

Power management and distribution

Figure F.13: Electronic devices temperature.

0 500 1000 1500

Time [min]

240

260

280

300

320

340

360

Tem

pera

ture

[K

]

Electronic devices temeprature

S-Band Transceiver

UHF Transceiver

SpaceCraft On-board computer

Payload On-board computer

Power management and distribution

Figure F.14: Electronic controllers temperature.

67

Chapter F. Thermal Control Subsystem Charts

0 500 1000 1500

Time [min]

270

275

280

285

290

295

300

305

310

315

320

Tem

pera

ture

[K

]

Battery temperature

Battery

Figure F.15: Battery temperature.

0 500 1000 1500

Time [min]

284

284.5

285

285.5

286

286.5

287

287.5

288

288.5

Tem

pera

ture

[K

]

Payload temperature

Payload

Figure F.16: Payload temperature.

68

G. Utilised Resources

• IDM-CIC: software developed by CNES to provide a technical reference during the concurrentengineering process required to design a satellite. Utilised for the concurrent design process of theCubeSat: mass budget, power consumption, power operational modes, temperature ranges, etc.• STK: Systems Tool Kit is a physics-based software package from AGI to perform complex analyses

of ground, sea, air, and space platforms. Utilised for the coverage analysis since it is a powerfultool that already implements the characteristics of G/S.• Stela: The Semi-analytic Tool for End of Life Analysis (STELA) is a software developed by CNES

utilised to perform the S/C de-orbit analysis.• GMAT: the General Mission Analysis Tool is a NASA open source software for space mission

design, optimisation, and navigation. Utilised for orbit propagation and orbit perturbation analysis.In addition, it has been used for eclipse calculation. Finally, it has been used for the validation ofSTELA reentry analysis.• Matlab: high-performance language for technical computing developed by MathWorks, utilised

for a wide range of applications: thermal model, link budget analysis, processing coverage datacoming from STK to set the overlap between two G/S coverage areas, etc.• Amsat Excel Sheet: utilised to validate the link budget computed with Matlab.• Microsoft Excel: utilised for a wide range of applications specially in the conceptual design phase:

power and mass estimations, link budget, etc.• Catia: software for computer-aided design developed by Dassault Systemes, utilised to compute

COTS CoG and Inertial properties for those not included in IDM-CIC library.• Fusion 360: software for computer-aided design developed by Autodesk. Utilised to create 3D

animations of the satellite.• LATEX: used to create the PDR report.• Google Drive: utilised for documents and scripts sharing. Use of Google Sheets to enhance parallel

work for certain common documents: requirements flow-down, risk analysis table, system datasummary, etc.• Slack: an instant messaging platform for groups developed by Slack Technologies, utilised for

interaction with other teams and NanoStar Challenge organisers.• Hangouts: utilised for team meetings.• Adobe Premiere: timeline-based video editing app developed by Adobe Systems. Utilised for

editing the project presentation.

H. System Data Summary

The System Data Summary is include in this section. This Management tool is utilised to ease theconcurrent design. For each selected COTS component, the main parameters are stored in an standardisedway.Only information coming from official manufacturer data is stored is added (not assumptions).Information is divided into different sections, so that the responsible for each subsystem has a clear viewof the relevant data for it. In order to store and update information unequivocally, there is a REF columnwhere the reference document from where information is taken is written. Besides, each field has a labelcomposed by DS-UNIT-XX (XX:number). There are some "empty fields" because the manufacturerdoes not provided that information. Nevertheless, they will be filled in future phases of the mission afterrequesting the information to such manufacturers. The different standard information sections are:

1. Reference Documents: datasheet name and version is stated here. Other reference document likeweb pages are also included.

2. General Information: manufacturer data, COTS model, TRL and number of units in the satelliteis stated

3. Mechanical Parameters: mass, CoG, inertia, mounting screws, etc.4. Thermal Parameters: qualification temperature ranges, thermo-optical properties, etc.5. Electrical Parameters: power consumption in different regimes, connectors6. Radiation Tolerance:7. Performances: main performance parameters of each component. This is the unique section that

depends on the component.

NanoStar Roscoff Worms Mission RW System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-RW-001

[RD-01] HT_RW210_Flyer DS-RW-002[RD-02]https://hyperiontechnologies.nl/products/rw210/ DS-RW-003[RD-03] DS-RW-004

2 General Information DS-RW-005Manufacturer HYPERION TECHNOLOGIES DS-RW-006Model RW210 SERIES 3.0 mNms DS-RW-007TRL 9 [RD-02] DS-RW-008Number of Units 4 DS-RW-009

3 Mechanical Parameters DS-RW-010Model [RD-01] Available upon request DS-RW-011Mass w/o margin g 32 [RD-01] In the table of the reference DS-RW-012Unit Margin [%] 5 CAT C: COTS DS-RW-013Mass g 33,6 Computed DS-RW-014Dimensions [x y z] mm [25 25 15] [RD-01] DS-RW-015Center of gravity [x y z] DS-RW-016Center of gravity ref frame DS-RW-017Moment of Inertia DS-RW-018Number of mounting screws 4 [RD-01] In the drawing DS-RW-019Type of mounting screws DS-RW-020CAD Model DS-RW-021Drawing DS-RW-022First Frequency Hz DS-RW-023Shock DS-RW-024Sine Load DS-RW-025Random Load DS-RW-026

4 Thermal Parameters DS-RW-027Power Dissipation Stand-By W DS-RW-028Power Dissipation ON W DS-RW-029Non-operation min temp °C DS-RW-030Non-operation max temp °C DS-RW-031Operation min temp °C -20 [RD-01] DS-RW-032Operation max temp °C 60 [RD-01] DS-RW-033Thermal Capacity J/KgK DS-RW-034Emissivity DS-RW-035Absortivity DS-RW-036Contact Conductivity W/m2K DS-RW-037Case Material DS-RW-038

5 Electrical Parameters DS-RW-039Voltage Input V 3,3 [RD-01] from 3.25V to 3.5V DS-RW-040Min Power Consumption W 0,01 [RD-01] In Idle DS-RW-041Nominal Power Consumption W 0,6 [RD-01] Nominal DS-RW-042Max Power Consumption W 1,5 [RD-01] Peak DS-RW-0435.1 Connectors DS-RW-044

Connector 1 16 pins [RD-01] DS-RW-045Connector 2 DS-RW-046Connector 3 DS-RW-047

6 Radiation Tolerance DS-RW-048Radiation Tolerance kRad >36 [RD-01] DS-RW-049

7 Performances DS-RW-050Maximum Torque mNm +/-0,1 [RD-01] DS-RW-051

NanoStar Roscoff Worms Mission RW System Data Sumary v1.0

Total Momentum Storage mNms +/-6,0 [RD-01] DS-RW-052Max Rotation Rate rpm 15000 [RD-01] DS-RW-053Control Accuracy [%] +/-0,5 [RD-01] % of target rpm DS-RW-054

NanoStar Roscoff Worms Mission STR System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-STR-001

[RD-01] HT_ST200_v2.1-flyer.pdf DS-STR-002[RD-02] https://hyperiontechnologies.nl/products/miniaturised-star-tracker/ DS-STR-003[RD-03] DS-STR-004

2 General Information DS-STR-005Manufacturer HYPERION TECHNOLOGIES DS-STR-006Model ST200 DS-STR-007TRL 9 [RD-02] DS-STR-008Number of Units 1 DS-STR-009

3 Mechanical Parameters DS-STR-010Model DS-STR-011Mass w/o margin g 42 [RD-01] Excluding baffle DS-STR-012Unit Margin [%] 5 [RD-01] DS-STR-013Mass g 44,1 [RD-01] DS-STR-014Dimensions [x y z] mm [29 29 38,1] [RD-01] X: optical axis; Y and Z outer bracket DS-STR-015Center of gravity [x y z] DS-STR-016Center of gravity ref frame DS-STR-017Moment of Inertia DS-STR-018Number of mounting screws 4 [RD-01] In the drawing DS-STR-019Type of mounting screws DS-STR-020CAD Model DS-STR-021First Frequency Hz DS-STR-022Drawing DS-STR-023Shock DS-STR-024Sine Load DS-STR-025Random Load DS-STR-026

4 Thermal Parameters DS-STR-027Power Dissipation Stand-By W DS-STR-028Power Dissipation ON W DS-STR-029Non-operation min temp °C DS-STR-030Non-operation max temp °C DS-STR-031Operation min temp °C -20 [RD-01] DS-STR-032Operation max temp °C 40 [RD-01] DS-STR-033Thermal Capacity J/KgK DS-STR-034Emissivity DS-STR-035Absortivity DS-STR-036Contact Conductivity W/m2K DS-STR-037Case Material DS-STR-038

5 Electrical Parameters DS-STR-039

Voltage Input V 3,65 [RD-01]Typical, maximum efficiency is reached when operating at the lowest voltage (3,6V)

DS-STR-040

Min Power Consumption W 0,18 [RD-01] DS-STR-041Nominal Power Consumption W 0,65 [RD-01] @3,65V and 5Hz of update rate DS-STR-042Max Power Consumption W 1 [RD-01] DS-STR-0435.1 Connectors DS-STR-044

Connector 1 pins 8 [RD-01] DS-STR-045Connector 2 DS-STR-046Connector 3 DS-STR-047

6 Radiation Tolerance DS-STR-048Radiation Tolerance kRad 9 [RD-01] DS-STR-049

7 Performances DS-STR-050Roll Attitude determination accuracy arcsec 200 [RD-01] arcsec (3sigma) DS-STR-051Pitch Attitude determination accuracy arcsec 30 [RD-01] arcsec (3sigma) DS-STR-052Yaw Attitude determination accuracy arcsec 30 [RD-01] arcsec (3sigma) DS-STR-053

NanoStar Roscoff Worms Mission STR System Data Sumary v1.0

Max slew rate (tip,tilt) °/s >0,3 [RD-01] DS-STR-054Max slew rate (roll) °/s >0,6 [RD-01] DS-STR-055Nominal Update rate Hz 5 [RD-01] DS-STR-056Data Output [RD-01] Digital and Interface protocol is RS422 DS-STR-057

NanoStar Roscoff Worms Mission SAS System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-SAS-001

[RD-01] nanoSSOC-D60-Technical-Specifications DS-SAS-002[RD-02] DS-SAS-003[RD-03] DS-SAS-004

2 General Information DS-SAS-005Manufacturer SOLAR MEMS DS-SAS-006Model nanoSSOC-D60 (2 orthogonal axes) DS-SAS-007TRL 9 [RD-01] orbiting since 2009 DS-SAS-008Number of Units 3 DS-SAS-009

3 Mechanical Parameters DS-SAS-010Model DS-SAS-011Mass w/o margin g 6,5 [RD-01] DS-SAS-012Unit Margin [%] 5 [RD-01] DS-SAS-013Mass g 6,825 [RD-01] DS-SAS-014Dimensions [x y z] mm [43 14 5,9] [RD-01] DS-SAS-015Center of gravity [x y z] DS-SAS-016Center of gravity ref frame DS-SAS-017Moment of Inertia DS-SAS-018Number of mounting screws 2 [RD-01] DS-SAS-019Type of mounting screws M2,5 [RD-01] DS-SAS-020CAD Model DS-SAS-021First Frequency Hz DS-SAS-022Drawing DS-SAS-023Shock g 3000g @ 1-100ms [RD-01] DS-SAS-024Sine Load DS-SAS-025Random Load g 14,1g @ 20-2000Hz [RD-01] DS-SAS-026

4 Thermal Parameters DS-SAS-027Power Dissipation Stand-By W 0,076 [RD-01] DS-SAS-028Power Dissipation ON W 0,076 [RD-01] DS-SAS-029Non-operation min temp °C DS-SAS-030Non-operation max temp °C DS-SAS-031Operation min temp °C -30 [RD-01] DS-SAS-032Operation max temp °C 85 [RD-01] DS-SAS-033Thermal Capacity J/KgK DS-SAS-034Emissivity 0,855 [RD-01] DS-SAS-035Absortivity 0,935 [RD-01] DS-SAS-036Contact Conductivity W/m2K DS-SAS-037Case Material [RD-01] Aluminum 6082; Black anodized DS-SAS-038

5 Electrical Parameters DS-SAS-039Voltage Input V 3,3 [RD-01] 5V under request DS-SAS-040Min Power Consumption W 0,07 [RD-01] 3,25V*21mA DS-SAS-041Nominal Power Consumption W 0,07 [RD-01] 3,3V*21mA DS-SAS-042Max Power Consumption W 0,08 [RD-01] 3,35*23mA DS-SAS-0435.1 Connectors DS-SAS-044

Connector 1: DF13A-10DP-1.25V(55) pins 10 [RD-01] From Hirose (female) DS-SAS-045Connector 2 DS-SAS-046Connector 3 DS-SAS-047

6 Radiation Tolerance DS-SAS-048Radiation Tolerance kRad 30 [RD-01] DS-SAS-049

7 Performances DS-SAS-050Field of View deg 60 [RD-01] Semiangle DS-SAS-051Accuracy in FOV deg 0,5 [RD-01] 3 sigma error DS-SAS-052Precision deg 0,1 [RD-01] DS-SAS-053

NanoStar Roscoff Worms Mission BAT System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-BAT-001

[RD-01] gs-ds-nanopower-bp4-27 DS-BAT-002[RD-02] gs-ds-nanopower-battery-17 DS-BAT-003[RD-03] DS-BAT-004

2 General Information DS-BAT-005Manufacturer GOMspace DS-BAT-006Model NanoPower BP4 DS-BAT-007TRL DS-BAT-008Number of Units 1 DS-BAT-009Cells Type [RD-01] Lithium Ion 18650 DS-BAT-010

3 Mechanical Parameters DS-BAT-011Model DS-BAT-012Mass w/o margin g 258 [RD-01] DS-BAT-013Unit Margin [%] 5 [RD-01] DS-BAT-014Mass g 270,9 [RD-01] DS-BAT-015Dimensions [x y z] mm [94 84 23] [RD-01] DS-BAT-016Center of gravity [x y z] DS-BAT-017Center of gravity ref frame DS-BAT-018Moment of Inertia DS-BAT-019Number of mounting screws 4 [RD-01] DS-BAT-020Type of mounting screws DS-BAT-021CAD Model DS-BAT-022First Frequency Hz DS-BAT-023Drawing [RD-01] Section 7 DS-BAT-024Shock g DS-BAT-025Sine Load DS-BAT-026Random Load g DS-BAT-027

4 Thermal Parameters DS-BAT-028Power Dissipation Stand-By W DS-BAT-029Power Dissipation ON W DS-BAT-030Non-operation min temp °C -20 [RD-02] Discharge DS-BAT-031Non-operation max temp °C 60 [RD-02] Discharge DS-BAT-032Operation min temp °C 0 [RD-02] Charge DS-BAT-033Operation max temp °C 45 [RD-02] Charge DS-BAT-034Thermal Capacity J/KgK DS-BAT-035Emissivity DS-BAT-036Absortivity DS-BAT-037Contact Conductivity W/m2K DS-BAT-038PCB Material [RD-01] Glass/Polymide IPC-A-610 cl. 3/A DS-BAT-039Heater Power W 6 [RD-01] Typical DS-BAT-040

5 Electrical Parameters DS-BAT-041Voltage Input V [RD-01] Depends on config. See section 2.2 of ref DS-BAT-042Min Power Consumption W DS-BAT-043Nominal Power Consumption W DS-BAT-044Max Power Consumption W DS-BAT-0455.1 Connectors DS-BAT-046

Connector 1: H2 Stack Connector pins 5 [RD-01] DS-BAT-047Connector 2: P1 Battery Connector pins 16 [RD-01] DS-BAT-048Connector 3: P2 Power-switch Output pins 4 [RD-01] Harwin M80-8670405 DS-BAT-049Connector 4: P3 Ground-break Connector pins 8 [RD-01] Harwin M80-8670805 DS-BAT-050

6 Radiation Tolerance DS-BAT-051Radiation Tolerance kRad DS-BAT-052

7 Performances DS-BAT-053Cycle Life cycles 350 [RD-02] 20% capacity loss: For 100% DOD and 25°C DS-BAT-054Battery capacity W/h 38,5 [RD-01] Table section 2.2 DS-BAT-055

NanoStar Roscoff Worms Mission SC OBC System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-OBC-001

[RD-01] gs-ds-nanomind-a3200_1006901.pdf DS-OBC-002[RD-02] DS-OBC-003[RD-03] DS-OBC-004

2 General Information DS-OBC-005Manufacturer GOMspace DS-OBC-006Model Nanomind A3200 DS-OBC-007TRL DS-OBC-008Number of Units 1 DS-OBC-009

3 Mechanical Parameters DS-OBC-010Model DS-OBC-011Mass w/o margin g 24 [RD-01] Including shield DS-OBC-012Unit Margin [%] 5 [RD-01] DS-OBC-013Mass g 25,2 [RD-01] DS-OBC-014Dimensions [x y z] mm [65 40 7,1] [RD-01] DS-OBC-015Center of gravity [x y z] DS-OBC-016Center of gravity ref frame DS-OBC-017Moment of Inertia DS-OBC-018Number of mounting screws 4 [RD-01] DS-OBC-019Type of mounting screws M3,2 [RD-01] DS-OBC-020CAD Model DS-OBC-021First Frequency Hz DS-OBC-022Drawing DS-OBC-023Shock DS-OBC-024Sine Load DS-OBC-025Random Load DS-OBC-026

4 Thermal Parameters DS-OBC-027Power Dissipation Stand-By W DS-OBC-028Power Dissipation ON W DS-OBC-029Non-operation min temp °C -30 [RD-01] DS-OBC-030Non-operation max temp °C 85 [RD-01] DS-OBC-031Operation min temp °C -30 [RD-01] DS-OBC-032Operation max temp °C 85 [RD-01] DS-OBC-033Thermal Capacity J/KgK DS-OBC-034Emissivity DS-OBC-035Absortivity DS-OBC-036Contact Conductivity W/m2K DS-OBC-037Case Material DS-OBC-038

5 Electrical Parameters DS-OBC-039Voltage Input V 3,2-3,4 [RD-01] DS-OBC-040Min Power Consumption W 0,01 [RD-01] In Idle DS-OBC-041Nominal Power Consumption W 0,17 [RD-01] Nominal DS-OBC-042Max Power Consumption W 0,9 [RD-01] Peak DS-OBC-0435.1 Connectors DS-OBC-044

Connector 1: P1 - Picoblade USART (debug) Connector pins 4 [RD-01] DS-OBC-045Connector 2: P2 - Picoblade Connector for JTAG pins 8 [RD-01] DS-OBC-046Connector 3: P6 - Picoblade Connector with PWM outputs pins 6 [RD-01] DS-OBC-047Connector 4: P7 - Picoblade Connector with I2C and VBAT pins 6 [RD-01] DS-OBC-048Connector 5: X1 - FSI Main Connectors pins 20 [RD-01] DS-OBC-049Connector 6: X3 - FSI Main Connectors pins 20 [RD-01] DS-OBC-050

6 Radiation Tolerance DS-OBC-051Radiation Tolerance kRad DS-OBC-052

7 Performances DS-OBC-053Processor [RD-01] AT32UC3C MCU: 32-bit RISC DS-OBC-054Processor clock Mhz 8-64 [RD-01] DS-OBC-055SDRAM MB 32 [RD-01] DS-OBC-056

NanoStar Roscoff Worms Mission PL OBC System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-FPGA-001

[RD-01] gs-ds-nanomind-z7000-15.pdf DS-FPGA-002[RD-02] DS-FPGA-003[RD-03] DS-FPGA-004

2 General Information DS-FPGA-005Manufacturer GOMspace DS-FPGA-006Model Nanomind Z700 DS-FPGA-007TRL DS-FPGA-008Number of Units 1 DS-FPGA-009

3 Mechanical Parameters DS-FPGA-010Model DS-FPGA-011Mass w/o margin g 76,8 [RD-01] Including shield (28,7g no shield) DS-FPGA-012Unit Margin [%] 5 [RD-01] DS-FPGA-013Mass g 80,64 [RD-01] DS-FPGA-014Dimensions [x y z] mm [65 40 6,5] [RD-01] DS-FPGA-015Center of gravity [x y z] DS-FPGA-016Center of gravity ref frame DS-FPGA-017Moment of Inertia DS-FPGA-018Number of mounting screws 4 [RD-01] DS-FPGA-019Type of mounting screws M3,2 [RD-01] DS-FPGA-020CAD Model DS-FPGA-021First Frequency Hz DS-FPGA-022Drawing DS-FPGA-023Shock DS-FPGA-024Sine Load DS-FPGA-025Random Load DS-FPGA-026

4 Thermal Parameters DS-FPGA-027Power Dissipation Stand-By W DS-FPGA-028Power Dissipation ON W DS-FPGA-029Non-operation min temp °C DS-FPGA-030Non-operation max temp °C DS-FPGA-031Operation min temp °C -40 [RD-01] DS-FPGA-032Operation max temp °C 85 [RD-01] DS-FPGA-033Thermal Capacity J/KgK DS-FPGA-034Emissivity DS-FPGA-035Absortivity DS-FPGA-036Contact Conductivity W/m2K DS-FPGA-037Case Material DS-FPGA-038

5 Electrical Parameters DS-FPGA-039Voltage Input V 3-3,6 [RD-01] DS-FPGA-040Min Power Consumption W 0,90 In Idle DS-FPGA-041Nominal Power Consumption W 2,30 Nominal DS-FPGA-042Max Power Consumption W 3,6 [RD-01] Peak DS-FPGA-0435.1 Connectors DS-FPGA-044

Connector 1 pins DS-FPGA-045Connector 2 pins DS-FPGA-046Connector 3 pins DS-FPGA-047

6 Radiation Tolerance DS-FPGA-048Radiation Tolerance kRad DS-FPGA-049

7 Performances DS-FPGA-050Processor [RD-01] ARM-based processor DS-FPGA-051Processor clock Mhz 800 [RD-01] DS-FPGA-052Mass Memory storage GB 32 [RD-01] DS-FPGA-053

NanoStar Roscoff Worms Mission PMAD System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-PMAD-001

[RD-01] gs-ds-nanopower-p60-dock-29 DS-PMAD-002[RD-02] DS-PMAD-003[RD-03] DS-PMAD-004

2 General Information DS-PMAD-005Manufacturer GOMSpace DS-PMAD-006Model Nanopower P60 with 2xPDU and 1xACU DS-PMAD-007TRL DS-PMAD-008Number of Units 1 DS-PMAD-009

3 Mechanical Parameters DS-PMAD-010Model DS-PMAD-011Mass w/o margin g 248 [RD-01] Taking 1 ACU and 2 PDU DS-PMAD-012Unit Margin [%] 5 [RD-01] DS-PMAD-013Mass g 260,4 [RD-01] DS-PMAD-014Dimensions [x y z] mm [92 88,9 28,8] [RD-01] Whole assembly DS-PMAD-015Center of gravity [x y z] DS-PMAD-016Center of gravity ref frame DS-PMAD-017Moment of Inertia DS-PMAD-018Number of mounting screws 4 [RD-01] DS-PMAD-019Type of mounting screws DS-PMAD-020CAD Model DS-PMAD-021First Frequency Hz DS-PMAD-022Drawing DS-PMAD-023Shock g DS-PMAD-024Sine Load DS-PMAD-025Random Load DS-PMAD-026

4 Thermal Parameters DS-PMAD-027Power Dissipation Stand-By W DS-PMAD-028Power Dissipation ON W DS-PMAD-029Non-operation min temp °C DS-PMAD-030Non-operation max temp °C DS-PMAD-031Operation min temp °C -35 [RD-01] DS-PMAD-032Operation max temp °C 85 [RD-01] DS-PMAD-033Thermal Capacity J/KgK DS-PMAD-034Emissivity DS-PMAD-035Absortivity DS-PMAD-036Contact Conductivity W/m2K DS-PMAD-037Case Material DS-PMAD-038

5 Electrical Parameters DS-PMAD-039Voltage Input V [RD-01] DS-PMAD-040Min Power Consumption W [RD-01] DS-PMAD-041Nominal Power Consumption W 0,6 [RD-01] ACU + 2xPDU DS-PMAD-042Max Power Consumption W DS-PMAD-0435.1 Connectors DS-PMAD-044

Connector 1 pins DS-PMAD-045Connector 2 pins DS-PMAD-046Connector 3 pins DS-PMAD-047

6 Radiation Tolerance DS-PMAD-048Radiation Tolerance kRad DS-PMAD-049

7 Performances DS-PMAD-050ACU inputs number 6 [RD-01] DS-PMAD-051PDU outputs number 9 [RD-01] DS-PMAD-052

NanoStar Roscoff Worms Mission Structure System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-ST-001

[RD-01] ISIS-CubeSat-Structures-Brochure-v2R-web_compressed.pdf DS-ST-002[RD-02] https://www.isispace.nl/product/6-unit-cubesat-structure/ DS-ST-003[RD-03] DS-ST-004

2 General Information DS-ST-005Manufacturer ISIS DS-ST-006Model 6-Unit CubeSat structure DS-ST-007TRL DS-ST-008Number of Units 1 DS-ST-009

3 Mechanical Parameters DS-ST-010Model Part avaialble (.stp) DS-ST-011Mass w/o margin g 1100 [RD-01] Primary + Secondary Mass DS-ST-012Unit Margin [%] 5 [RD-01] DS-ST-013Mass g 1155 [RD-01] DS-ST-014Dimensions [x y z] mm [100 226,3 340,5] [RD-01] DS-ST-015Center of gravity [x y z] DS-ST-016Center of gravity ref frame DS-ST-017Moment of Inertia DS-ST-018Number of mounting screws DS-ST-019Type of mounting screws DS-ST-020CAD Model DS-ST-021First Frequency Hz DS-ST-022Drawing DS-ST-023Shock g DS-ST-024Sine Load DS-ST-025Random Load g DS-ST-026

4 Thermal Parameters DS-ST-027Power Dissipation Stand-By W DS-ST-028Power Dissipation ON W DS-ST-029Non-operation min temp °C DS-ST-030Non-operation max temp °C DS-ST-031Operation min temp °C -40 [RD-02] DS-ST-032Operation max temp °C 80 [RD-02] DS-ST-033Thermal Capacity J/KgK DS-ST-034Emissivity DS-ST-035Absortivity DS-ST-036Contact Conductivity W/m2K DS-ST-037Case Material DS-ST-038

5 Radiation Tolerance DS-ST-039Radiation Tolerance kRad DS-ST-040

6 Performances DS-ST-041

NanoStar Roscoff Worms Mission Magnetorquer System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-MGNT-001

[RD-01] iMTQ-Brochure-v1.pdf DS-MGNT-002[RD-02] https://www.isispace.nl/product/isis-magnetorquer-board/ DS-MGNT-003[RD-03] DS-MGNT-004

2 General Information DS-MGNT-005Manufacturer ISIS DS-MGNT-006Model ISIS Magnetorquer Board DS-MGNT-007TRL Used in MarCO DS-MGNT-008Number of Units 1 DS-MGNT-009

3 Mechanical Parameters DS-MGNT-010Model Part avaialble (.stp) DS-MGNT-011Mass w/o margin g 196 [RD-01] DS-MGNT-012Unit Margin [%] 5 [RD-01] DS-MGNT-013Mass g 205,8 [RD-01] DS-MGNT-014Dimensions [x y z] mm [95,9 90,1 17] [RD-01] DS-MGNT-015Center of gravity [x y z] DS-MGNT-016Center of gravity ref frame DS-MGNT-017Moment of Inertia DS-MGNT-018Number of mounting screws DS-MGNT-019Type of mounting screws DS-MGNT-020CAD Model DS-MGNT-021First Frequency Hz DS-MGNT-022Drawing DS-MGNT-023Shock g DS-MGNT-024Sine Load DS-MGNT-025Random Load g DS-MGNT-026

4 Thermal Parameters DS-MGNT-027Power Dissipation Stand-By W DS-MGNT-028Power Dissipation ON W DS-MGNT-029Non-operation min temp °C DS-MGNT-030Non-operation max temp °C DS-MGNT-031Operation min temp °C -40 [RD-01] DS-MGNT-032Operation max temp °C 70 [RD-01] DS-MGNT-033Thermal Capacity J/KgK DS-MGNT-034Emissivity DS-MGNT-035Absortivity DS-MGNT-036Contact Conductivity W/m2K DS-MGNT-037Case Material DS-MGNT-038

5 Electrical Parameters DS-MGNT-039Voltage Input V 5 [RD-01] DS-MGNT-040Min Power Consumption W 0,175 [RD-01] No actuation DS-MGNT-041Nominal Power Consumption W 1,2 DS-MGNT-042Max Power Consumption W <1,2 [RD-01] Full actuation (3-Axis) DS-MGNT-0435.1 Connectors DS-MGNT-044

Connector 1 pins DS-MGNT-045Connector 2 pins DS-MGNT-046Connector 3 pins DS-MGNT-047

6 Radiation Tolerance DS-MGNT-048Radiation Tolerance kRad DS-MGNT-049

7 Performances DS-MGNT-050Detumbling frequency Hz 1-8 DS-MGNT-051

NanoStar Roscoff Worms Mission S-Band Transceiver System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-SBT-001

[RD-01] gs-ds-nanocom-tr600-16.pdf DS-SBT-002[RD-02] DS-SBT-003[RD-03] DS-SBT-004

2 General Information DS-SBT-005Manufacturer GOMspace DS-SBT-006Model NanoCom TR-600 DS-SBT-007TRL DS-SBT-008Number of Units 1 DS-SBT-009

3 Mechanical Parameters DS-SBT-010Model DS-SBT-011Mass w/o margin g 65,3 [RD-01] Mass total PCB + shield DS-SBT-012Unit Margin [%] 5 [RD-01] DS-SBT-013Mass g 68,565 [RD-01] DS-SBT-014Dimensions [x y z] mm [65 40 14,8] [RD-01] DS-SBT-015Center of gravity [x y z] DS-SBT-016Center of gravity ref frame DS-SBT-017Moment of Inertia DS-SBT-018Number of mounting screws DS-SBT-019Type of mounting screws DS-SBT-020CAD Model DS-SBT-021First Frequency Hz DS-SBT-022Drawing DS-SBT-023Shock g DS-SBT-024Sine Load DS-SBT-025Random Load g DS-SBT-026

4 Thermal Parameters DS-SBT-027Power Dissipation Stand-By W DS-SBT-028Power Dissipation ON W DS-SBT-029Non-operation min temp °C DS-SBT-030Non-operation max temp °C DS-SBT-031Operation min temp °C -40 [RD-01] DS-SBT-032Operation max temp °C 85 [RD-01] DS-SBT-033Thermal Capacity J/KgK DS-SBT-034Emissivity DS-SBT-035Absortivity DS-SBT-036Contact Conductivity W/m2K DS-SBT-037Case Material DS-SBT-038

5 Electrical Parameters DS-SBT-039Voltage Input V 3 to 3,6 [RD-01] DS-SBT-040Min Power Consumption W 3,88 [RD-01] Idle DS-SBT-041Nominal Power Consumption W 4,1 [RD-01] Nominal DS-SBT-042Max Power Consumption W DS-SBT-0435.1 Connectors DS-SBT-044

Connector 1: J1 – Rx 1 pins 2 [RD-01] DS-SBT-045Connector 2: J2 – Tx 1 pins 2 [RD-01] DS-SBT-046Connector 3: J3 – Rx 2 pins 2 [RD-01] DS-SBT-047Connector 4: J3 – Rx 2 pins 2 [RD-01] DS-SBT-048

6 Radiation Tolerance DS-SBT-049Radiation Tolerance kRad DS-SBT-050

7 Performances DS-SBT-051PCB Material [RD-01] Glass/polyimide ESA ECSS-Q-ST-70-11-C DS-SBT-052Band 70 MHz to 6 GHz [RD-01] DS-SBT-053Noise Figure dB 7,1-6,1 [RD-01] @20ºC and 55-70 Rx Gain Index DS-SBT-054

NanoStar Roscoff Worms Mission S-Band Antenna System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-SBA-001

[RD-01] gs-ds-nanocom-ant2000.pdf DS-SBA-002[RD-02] DS-SBA-003[RD-03] DS-SBA-004

2 General Information DS-SBA-005Manufacturer GOMspace DS-SBA-006Model NanoCom ANT2000 DS-SBA-007TRL DS-SBA-008Number of Units 1 DS-SBA-009

3 Mechanical Parameters DS-SBA-010Model DS-SBA-011Mass w/o margin g 110 [RD-01] Approximate, depends on mounting, etc. DS-SBA-012Unit Margin [%] 5 [RD-01] DS-SBA-013Mass g 115,5 [RD-01] DS-SBA-014Dimensions [x y z] mm [98 98 20,1] [RD-01] DS-SBA-015Center of gravity [x y z] DS-SBA-016Center of gravity ref frame DS-SBA-017Moment of Inertia DS-SBA-018Number of mounting screws DS-SBA-019Type of mounting screws DS-SBA-020CAD Model DS-SBA-021First Frequency Hz DS-SBA-022Drawing DS-SBA-023Shock g DS-SBA-024Sine Load DS-SBA-025Random Load g DS-SBA-026

4 Thermal Parameters DS-SBA-027Power Dissipation Stand-By W DS-SBA-028Power Dissipation ON W DS-SBA-029Non-operation min temp °C -40 [RD-01] DS-SBA-030Non-operation max temp °C 85 [RD-01] DS-SBA-031Operation min temp °C -40 [RD-01] DS-SBA-032Operation max temp °C 85 [RD-01] DS-SBA-033Thermal Capacity J/KgK DS-SBA-034Emissivity DS-SBA-035Absortivity DS-SBA-036Contact Conductivity W/m2K DS-SBA-037Case Material DS-SBA-038

5 Electrical Parameters DS-SBA-039Voltage Input V 8 to 18 [RD-01] DS-SBA-040Min Power Consumption W DS-SBA-041Nominal Power Consumption W 10,7 [RD-01] DS-SBA-042Max Power Consumption W 11,0 [RD-01] DS-SBA-0435.1 Connectors DS-SBA-044

Connector 1: J102 - RX RF COAXIAL CONNECTORpins 2 [RD-01] DS-SBA-045Connector 2: J300 - TX RF COAXIAL CONNECTORpins 2 [RD-01] DS-SBA-046Connector 3: J400 - Power Connector pins 7 [RD-01] DS-SBA-047Connector 4: J401 - Control Connector pins 10 [RD-01] DS-SBA-048Connector 5: J402 - Debug pins 6 [RD-01] DS-SBA-049

6 Radiation Tolerance DS-SBA-050Radiation Tolerance kRad DS-SBA-051

7 Performances DS-SBA-052

NanoStar Roscoff Worms Mission UHF Transceiver System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-UHFT-001

[RD-01] gs-ds-nanocom-ant2000.pdf DS-UHFT-002[RD-02] DS-UHFT-003[RD-03] DS-UHFT-004

2 General Information DS-UHFT-005Manufacturer GOMspace DS-UHFT-006Model NanoCom AX100 DS-UHFT-007TRL DS-UHFT-008Number of Units 1 DS-UHFT-009

3 Mechanical Parameters DS-UHFT-010Model DS-UHFT-011Mass w/o margin g 24,5 [RD-01] DS-UHFT-012Unit Margin [%] 5 [RD-01] DS-UHFT-013Mass g 25,725 [RD-01] DS-UHFT-014Dimensions [x y z] mm [65 40 6,5] [RD-01] DS-UHFT-015Center of gravity [x y z] DS-UHFT-016Center of gravity ref frame DS-UHFT-017Moment of Inertia DS-UHFT-018Number of mounting screws DS-UHFT-019Type of mounting screws DS-UHFT-020CAD Model DS-UHFT-021First Frequency Hz DS-UHFT-022Drawing DS-UHFT-023Shock g DS-UHFT-024Sine Load DS-UHFT-025Random Load g DS-UHFT-026

4 Thermal Parameters DS-UHFT-027Power Dissipation Stand-By W DS-UHFT-028Power Dissipation ON W DS-UHFT-029Non-operation min temp °C -30 [RD-01] DS-UHFT-030Non-operation max temp °C 85 [RD-01] DS-UHFT-031Operation min temp °C -30 [RD-01] DS-UHFT-032Operation max temp °C 85 [RD-01] DS-UHFT-033Thermal Capacity J/KgK DS-UHFT-034Emissivity DS-UHFT-035Absortivity DS-UHFT-036Contact Conductivity W/m2K DS-UHFT-037Case Material DS-UHFT-038

5 Electrical Parameters DS-UHFT-039Voltage Input V 3,3-3,4 [RD-01] DS-UHFT-040Min Power Consumption W 0,18 [RD-01] DS-UHFT-041Nominal Power Consumption W 2,6 [RD-01] DS-UHFT-042Max Power Consumption W 3,3 [RD-01] DS-UHFT-0435.1 Connectors DS-UHFT-044

Connector 1: J2 - Picoblade Connector pins 8 [RD-01] DS-UHFT-045Connector 2: J3 - Picoblade USART (debug) Connectorpins 4 [RD-01] DS-UHFT-046Connector 3: J4 - MCX RF Connector pins 7 [RD-01] DS-UHFT-047Connector 4: J1 - FSI Main Connector pins 20 [RD-01] DS-UHFT-048

11 DS-UHFT-049Radiation Tolerance kRad DS-UHFT-050

7 Performances DS-UHFT-051Frequency bands [RD-01] UHF and VHF DS-UHFT-052Noise Figure dB ~1 [RD-01] DS-UHFT-053Data rate kbps 0,1-38,4 [RD-01] DS-UHFT-054Sensitivity dBm -137 [RD-01] DS-UHFT-055

NanoStar Roscoff Worms Mission UHF Antenna System Data Sumary v1.0

FIELD UNITS VALUE REF COMMENTS ID1 Reference Documents DS-UHFA-001

[RD-01] gs-ds-nanocom-ant6f-uhf-21.pdf DS-UHFA-002[RD-02] DS-UHFA-003[RD-03] DS-UHFA-004

2 General Information DS-UHFA-005Manufacturer GOMspace DS-UHFA-006Model NanoCom ANT-6F UHF DS-UHFA-007TRL DS-UHFA-008Number of Units 1 DS-UHFA-009

3 Mechanical Parameters DS-UHFA-010Model DS-UHFA-011Mass w/o margin g 90 [RD-01] DS-UHFA-012Unit Margin [%] 5 [RD-01] DS-UHFA-013Mass g 94,5 [RD-01] DS-UHFA-014Dimensions [x y z] mm [221,7 116,7 5,3] [RD-01] DS-UHFA-015Center of gravity [x y z] DS-UHFA-016Center of gravity ref frame DS-UHFA-017Moment of Inertia DS-UHFA-018Number of mounting screws DS-UHFA-019Type of mounting screws DS-UHFA-020CAD Model DS-UHFA-021First Frequency Hz DS-UHFA-022Drawing DS-UHFA-023Shock g DS-UHFA-024Sine Load DS-UHFA-025Random Load g DS-UHFA-026

4 Thermal Parameters DS-UHFA-027Power Dissipation Stand-By W DS-UHFA-028Power Dissipation ON W DS-UHFA-029Non-operation min temp °C DS-UHFA-030Non-operation max temp °C DS-UHFA-031Operation min temp °C -40 [RD-01] DS-UHFA-032Operation max temp °C 85 [RD-01] DS-UHFA-033Thermal Capacity J/KgK DS-UHFA-034Emissivity DS-UHFA-035Absortivity DS-UHFA-036Contact Conductivity W/m2K DS-UHFA-037Case Material DS-UHFA-038

5 Electrical Parameters DS-UHFA-039Voltage Input V 3,2-3,4 [RD-01] DS-UHFA-040Min Power Consumption W DS-UHFA-041Nominal Power Consumption W 0,2 [RD-01] DS-UHFA-042Max Power Consumption W DS-UHFA-0435.1 Connectors DS-UHFA-044

Connector 1: P1 – External Programmer pins [RD-01] DS-UHFA-045Connector 2: P2 – GomSpace Release Bus (GSRB) Connectorpins 8 [RD-01] DS-UHFA-046Connector 3: P3 – GomSpace Release Bus (GSRB) Connectorpins 8 [RD-01] DS-UHFA-047Connector 4: J3 - SMPM RF Connector pins [RD-01] DS-UHFA-048Connector 5: J4 - SMPM RF Connector pins [RD-01] DS-UHFA-049Connector 6: J5 - MCX RF Connector pins [RD-01] DS-UHFA-050

11 DS-UHFA-051Radiation Tolerance kRad DS-UHFA-052

7 Performances DS-UHFA-053Frequency range MHz 340-680 [RD-01] DS-UHFA-054Others [RD-01] Integrated antenna release system DS-UHFA-055