team bernoulli report - nanostarv 1.0 1 team bernoulli report gonzalo azaña, upm tomás girona, upm...

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Team Bernoulli report Gonzalo Azaña, UPM

Tomás Girona, UPM

Daniel Gómez, UPM

Néstor Martínez, UPM

José Ramírez, UPM

NANOSTAR consortium

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Project funded by the Interreg Sudoe Programme through the European Regional Development Fund (ERDF)

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TABLE OF CONTENTS

EXECUTIVE ABSTRACT ................................................................................................. 4

LIST OF FIGURES ........................................................................................................... 5

LIST OF TABLES ............................................................................................................. 5

INTRODUCTION ............................................................................................................. 7

Scope ...................................................................................................................................................... 7

Team composition ................................................................................................................................. 7

APPLICABLE AND REFERENCE DOCUMENTS ........................................................... 8

References ............................................................................................................................................. 8

Change Record Log ................................................................................................................................ 9

MISSION OVERVIEW, REQUIREMENTS FLOWDOWN AND TESTS ......................... 11

Mission Objectives and Requirements ................................................................................................ 11

Team Bernoulli Solution ....................................................................................................................... 12

Management Plan ................................................................................................................................ 12

Team Management & Communication ............................................................................................... 13

Methodology and resources ................................................................................................................ 13

SUBSYSTEMS ANALISYS AND DESIGN ...................................................................... 15

Mission analysis (MA) .......................................................................................................................... 15

Payload (PL) .......................................................................................................................................... 18

System operations modes.................................................................................................................... 18

Space propulsion subsystems (SPS) ..................................................................................................... 19

Attitude, determination and control subsystem (ADCS) ..................................................................... 19

Command and data handling (C&DH).................................................................................................. 24

Communication subsystem (CS) and ground segment (GS) ................................................................ 24

Electric power subsystem (EPS) ........................................................................................................... 27

Mechanical design and structure (MD&S) ........................................................................................... 32

Structure ........................................................................................................................................... 32

Configuration .................................................................................................................................... 33

Deployment of solar cells ................................................................................................................. 35

Mass budget ..................................................................................................................................... 35

Thermal control subsystem (TCS) ........................................................................................................ 37

Launcher ............................................................................................................................................... 42

RISK ANALYSIS AND MITIGATION............................................................................. 43

Risk policy ............................................................................................................................................. 43

Risk strategy ..................................................................................................................................... 43

Risk severity scoring ......................................................................................................................... 44

Risk likelihood scoring ...................................................................................................................... 44

Risk index and magnitude scheme ................................................................................................... 45

Acceptance criteria and actions proposed ....................................................................................... 45

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Risk Assesment ..................................................................................................................................... 46

Method of assesment ....................................................................................................................... 46

Principle ............................................................................................................................................ 46

Identified risk scenarios and assesment ........................................................................................... 46

Risk trend .......................................................................................................................................... 49

CONCLUSIONS AND FUTURE WORK ........................................................................50

APPENDICES ................................................................................................................. 51

Appendix I: Glossary of Acronyms ....................................................................................................... 52

Appendix II: GANTT diagram ................................................................................................................ 53

Appendix III: ESA Network ................................................................................................................... 54

appendix IV: Risk registers ................................................................................................................... 55

Appendix V: Bottom-Up cost estimation ............................................................................................. 60

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EXECUTIVE ABSTRACT

The mission proposed by the NanoStar consortium is the designing of a CubeSat in order to study behavior in space conditions of the Roscoff worms. These animals can be a powerful ally in future manned missions, due to its capacity of producing O2 and recycle the CO2.

This document outlines the principal points about the management and technical processes followed during the whole challenge in order to achieve an engineering solution capable of performing the mission and the requirements established.

The whole management plan is explained, from the scheduled designed and the miscellaneous to be achieved; to the team management, where each of the members in the Bernoulli Team have performed a role during the whole project.

Moreover, the technical design has been divided in different subsystems, where each of the members was responsible of two of them. The solution reached is a 6U satellite capable of perform the scientific mission and transmit the data to an ESA ground stations for 3 months and even more if it was necessary. All the technical requirements by the NanoStar consortium have been accomplished. Moreover, the components used are COTS with a high TRL and state-of-the-art technology.

The risk policy has been defined too. A risk analysis has been done in order to mitigate the severity in the mission performance of different scenarios, applying the ECSS standards.

Finally, some improvements are stated in the design but lowering the TRL of the components, for possible future missions with the same objective but enhancing the performance of the payload. This means a better quality of the scientific data to be analyzed.

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LIST OF FIGURES

Figure 1: Organigram of the team .................................................................................. 7

Figure 2: Shape of the designed orbit .......................................................................... 15

Figure 3: Deorbitation time for a circular orbit of altitude 567 km ......................... 17

Figure 4:Deorbitation time for a circular orbit of altitude 410 km ........................... 17

Figure 5: Angles and distance considered for the CS analysis. .............................. 25

Figure 6: Diagram of the electrical power circuit...................................................... 32

Figure 7:6U structure ...................................................................................................... 33

Figure 8: Detailed view of the components ............................................................... 34

Figure 9: DEPLOYMENT MECHANISM ........................................................................ 35

Figure 10: Centre of gravity and inertia of the satellite ............................................ 37

Figure 11: Scheme of principal axes and radiation origin ......................................... 38

Figure 12: Definition of the internal arrangement ..................................................... 40

Figure 13: Scheme of the radiation absortion of the payload module ................... 41

Figure 14: Payload charge to sunsynchronous orbit for rocket lab electron launcher ............................................................................................................................ 42

Figure 15: Deployment margins for the launcher ...................................................... 42

Figure 16: Global ESA network. (source (5)) ................................................................ 54

Figure 17: Breakdown structure of the mission ......................................................... 60

Figure 18: Bottom-Up cost estimation ......................................................................... 61

LIST OF TABLES

Table 1: Change record performed ............................................................................... 10

Table 2: Top level mission requirements. ..................................................................... 11

Table 3: Sowftware tool used for subsystem design ................................................ 13

Table 4: Limit orbits for the mission ............................................................................. 16

Table 5: PL data generation ........................................................................................... 18

Table 6: COntrol modes of the Satellite ...................................................................... 20

Table 7:Solar radiation pressure torque coefficients ................................................ 20

Table 8: Aerodynamic torque coefficients .................................................................. 21

Table 9: Gravity-gradient torque coefficients ............................................................ 21

Table 10: Magnetic torque coefficients ........................................................................ 21

Table 11: Estimation of disturbance torque ................................................................. 21

Table 12: Performance of the CubeWheel reaction wheels .................................... 23

Table 13: Carrier to noise power density for different orbital altitudes................. 26

Table 14: Time frame with Kiruna ground stations for different orbital altitudes 27

Table 15: Data rate for each of the orbits .................................................................... 27

Table 16: Eb/No for each of the orbits .......................................................................... 27

Table 17: Electtrical power budget .............................................................................. 28

Table 18: Mass budget .................................................................................................... 36

Table 19: Radiation sources ........................................................................................... 37

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Table 20:Dissipation of internal components ............................................................ 38

Table 21: Possible materials for coating the spacecraft ........................................... 40

Table 22: Temperature definition for the Hot and Cold cases ................................ 40

Table 23: Definition of the hot and cold case ............................................................. 41

Table 24: Hot and cold case for the payload module ................................................ 41

Table 25: Severity scale of risks .................................................................................... 44

Table 26: Likelihood ocurrence of risks ....................................................................... 44

Table 27: Risk index scheme ......................................................................................... 45

Table 28: Individual risk action criteria ........................................................................ 45

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INTRODUCTION

SCOPE

The principal objective of the project is to perform the design of a cubesat satellite from phase 0 to phase A.

The objectives of this report are:

• Present the project management followed during the challenge. • Present the solution proposed to the challenge and the design process

behind. • Present the risks identified and their mitigation.

TEAM COMPOSITION

The team was created in September by 5 members of the MSc in Space System Engineering from the UPM. Each chief subsystem role was issued depending on the background, tool handling and preference of each member. Moreover, in order to facilitate the management of the project, two team leaders were selected: a system engineer, responsible of the technical part which has a professional background in satellite design; and a project manager, responsible of supervising the correct development and planning of the project which also

has professional background in planning engineering. The organigram of the team is clearly showed in Figure 1.

The creation of the deliverables will be a cooperative work of every member of the team, except the IDM-CIC file which will be made by the MD&S chief. However, the responsibility of the acceptance and content of every deliverable will be always the System Engineer and the Project Manager.

FIGURE 1: ORGANIGRAM OF THE TEAM

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APPLICABLE AND REFERENCE DOCUMENTS

REFERENCES

1. NANOSTAR project. Preliminary Design. [Online] 2019. [Cited: 12 14, 2019.] https://nanostar-project.gitlab.io/main/source/preliminary-design/index.html.

2. Wertz, James R., Everett, David F. and Puschell, Jeffery J. Space Mission Engineering: The New SMAD. Hawthorne : Microcosm Press, 2015. ISBN 978-1-881-883-15-9.

3. Cubespace. CubeWheel. [Online] Cubespace, 2019. [Cited: November 28, 2019.] https://cubespace.co.za/cubewheel/.

4. QB50. QB50. QB50. [Online] EU. https://www.qb50.eu/index.php/tech-docs/category/QB50_Systems_Requirements_issue_76e8e.pdf?download=89:qb50-docs.

5. ISIS. ISIS On board computer. [Online] [Cited: December 29, 2019.] https://www.isispace.nl/wp-content/uploads/2016/02/IOBC-Brochure-web-compressed.pdf.

6. NASA. State of the art. Small spacecraft technology. Moffet Field : NASA Center for AeroSpace Information, 2018. NASA/TP—2018–220027.

7. ESA. Estrack ground stations. [Online] European Space Agency, 2019. [Cited: 12 14, 2019.] https://www.esa.int/Enabling_Support/Operations/Estrack/Estrack_ground_stations.

8. ENDUROSAT. CubeSat communication mocules. [Online] 2019. [Cited: December 3, 2019.] https://www.endurosat.com/products/#cubesat-communication-modules.

9. —. [Online] [Cited: November 29, 2019.] https://www.endurosat.com/cubesat-store/all-cubesat-modules/s-band-antenna-commercial/.

10. GOMSPACE. Datasheet Nanopower P110. [Online] October 22, 2018. [Cited: December 6, 2019.] https://gomspace.com/UserFiles/Subsystems/datasheet/gs-ds-nanopower-p110-210.pdf.

11. ECSS secretariat. Space Engineering Space Environment. Noordwijk : ECSS secretariat, 2008. ECSS-E-ST-10-04C.

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12. Rocket lab. Payload user's guide. [Online] June 2019. [Cited: December 15, 2019.] https://www.rocketlabusa.com/assets/Uploads/Rocket-Lab-Payload-Users-Guide-6.4.pdf.

13. NANOSTAR Consortium. Preliminary design challenge. Mission requirements. [Online] December 5, 2019. [Cited: December 6, 2019.] https://nanostarproject.eu/wp-content/uploads/2019/12/NANOST-REQ-070-High-level-space-mission-requirements-for-Phase-1-Challenge-2nd-Edition.pdf.

14. ECSS Secretariat. ECSS-M-ST-80C . Noordwijk : ECSS, 2008.

15. GOMSPACE. Datasheet S-Band active antenna. [Online] 2019. [Cited: December 5, 2019.] https://gomspace.com/UserFiles/Subsystems/datasheet/gs-ds-nanocom-ant2000.pdf.

CHANGE RECORD LOG

Edition/Revision Date Description of the change Reponsible

V1.0 13/12/2019 Creation of the document: • Added Scope section • Added Team Composition

section • Added Management plan

section • Added Team Management

section • Added Methodology and

Resources section • Added Risk Policy section • Added Appendix I • Added Appendix II

Tomás Girona

V1.1 14/12/2019 • Updated Risk Policy section

• Added Team Bernoulli Solution section

Tomás Girona

V 1.2 27/12/2019 • Updated Subsystem analysis and design section (ADCS and TCS parts)

Jose Luis Ramirez

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V 2.0 28/12/2019 • Updated and adding information to all the technical subsections

Whole team

V 2.1 29/12/2019 • Check of the PM and the SE

Tomás Girona Daniel Gómez

TABLE 1: Change record performed

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MISSION OVERVIEW, REQUIREMENTS FLOWDOWN AND TESTS

MISSION OBJECTIVES AND REQUIREMENTS

The mission objective is to create a demonstrator capable of maintain alive and retrieve information of a colony of Roskoff worms in a space environment. For this purpose, a CubeSat standard satellite is going to be designed.

Other CubeSat missions with biological payload has been GeneSat(2006), PharmaSat(, O/OREOS(2010), SporeSat (2014), all of them with a 3U configuration; or EcAMSat(2017) with a configuration of 6U and a payload of 3U, similar to the configuration of NanoStar challenge.

The top-level mission requirements demanded by the client are,

Item Requirement RW-1 The system shall carry, activate and operate safely the main scientific

payload RW-2 scientific data obtained by the main science payload shall be

transmitted to Earth within the mission timeframe. RW-3 The satellite shall be capable of performing the mission objectives,

considering the space environment constraints. RW-4 The satellite shall guarantee the correct attitude RW-5 The satellite volume shall not exceed 8U RW-6 The mission duration from launch to end-of-life shall not be lower than

2 weeks and ideally up to 3 months. RW-7 The maximum mission duration from launch to end-of-life shall comply

with space regulation RW-8 Ground segment(s) shall rely only on the tracking stations of the ESA

network RW-9 During all its lifecycle, the nanosatellite shall not introduce any of its

embedded photosymbiotic flat wormsor its associated viruses or bacteria population into any area on Earth

RW-10

The conception of the nanosatellite shall ensure that all its lifecycle tooling and environment is ROHS, REACH and CSR compliance.

TABLE 2: TOP LEVEL MISSION REQUIREMENTS.

All the requirements require tracking during the whole project in order to know their compliment. Moreover, in order to know the feasibility of the requirements, COTS components with high TRL are going to be selected. Furthermore, the preference of the SC will be to use components with fly-heritage traceable, in order to assure an extreme high likelihood of requirements accomplishement.

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TEAM BERNOULLI SOLUTION

The solution proposed is a 6U satellite with 6U solar surface panels facing directly to the sun during the whole mission, and 2U solar surfaces panels to be deployed if needed. The Sunsynchronous orbit designed allows the correct performance of the satellite during the whole mission without the need of heaters or extra batteries (which are needed during the launching), favorizing the design of a very simple, economic and efficient spacecraft.

All the subsystems components are COTS with a high TRL, meaning a very trustable performance and a cheaper design.

In order to achieve a consistent solution, the components are COTS with a TRL-8/9, lowing the risk likelihood, the total costs and the feasibility of the solution.

MANAGEMENT PLAN

The project has been divided in three phases:

Phase 1:

• Role, schedule and communication protocols definition. • Risk policy definition. • Revision of the mission requirements and definition of the rest of technical

requirements.

This phase finishes with the updating of the mission requirements after the conference call taking place the 14th of November.

Phase 2:

• Design of the satellite following a concurrent design methodology. • Risk identification, mitigation and acceptance

This phase finishes with the compliance of the system engineer for all the solutions of the subsystems chiefs and the final acceptance of every risk identified and mitigated.

Phase 3:

• Creation of the report with the solution proposed, the project management and the risk analysis.

• Creation of the IDM-SIS deliverable. • Creation of the video explaining the solution.

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Each of the deliverables will be considered completed once the project manager and the system engineer have checked and accepted them.

Submission and end of NanoStar challenge

Christmas holiday has been an issue in order to perform presential sessions. Therefore, most part of design process has been arranged before the 20th of December, so the concurrent design sessions could be performed with the whole team. Moreover, during the development of the planning, delays have been considered exaggerating the estimated time of all the phases. A GANTT diagram was created for an easier visualization of the project schedule and planned phases (see Appendix II: GANTT diagram).

TEAM MANAGEMENT & COMMUNICATION

The team management has been designed considering holidays, viability of the members, access to software and phases of the project. From these considerations the following protocols have been raised:

• Two hours presential meeting twice a week (Monday and Wednesday) from the beginning of the project until the beginning of Christmas Holidays.

• Two hours on-line meeting twice a week (Monday and Wednesday) from the beginning of the Christmas Holidays until the end of the project.

• On-site availability of the PM and the SE every Friday morning if needed until the beginning of Christmas Holidays.

• Telematic availability of the PM and the SE during the whole weekend. • Use of Google Meet for telematic meetings. • Use of a WhatsApp group for instant messaging between members.

METHODOLOGY AND RESOURCES

A Concurrent Design methodology has been implemented for the design of the spacecraft’s subsystems. Each subsystem chief has followed the preliminary design methodology provided by NANOSTAR consortium (1) and the SMAD (2), being the principal references for the preliminary design of the satellite.

Moreover, some subsystems have used specialized software tools in order to perform the design (see Table 3).

Mission Analysis STK, GMAT Electric power IDM-CIC Mechanical Design CATIA V5, IDM-CIC

TABLE 3: SOWFTWARE TOOL USED FOR SUBSYSTEM DESIGN

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In order to correctly perform a concurrent design, the SE has updated data of all subsystems so tracking and checking can be made instantly. For this, the team has implemented a shared google Spreadsheet, an easy to use tool were design calculations and tracking can be made. Tracking of the report versions has been made with Microsoft Word track changes tool, so both the PM and the SE can view all the changes made to the document.

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SUBSYSTEMS ANALISYS AND DESIGN

MISSION ANALYSIS (MA)

The mission will take place in a Repeating-Sunsynchronus down-dusk circular orbit. Sunsynchronus down-dusk is a high inclination orbit which has long periods without eclipses. The mission launching can be design in order to have a non-eclipse period longer than the mission lifetime. It is a very common destination of other missions because this orbit reduces the severity of a possible battery fail.

Repeating Ground Track (RGT) orbit main characteristic is its period, which is a fraction of the Earth one. This means that the satellite passes over the same places each k cycles. This way is easier to determinate the access over the ground stations.

The design of the orbit is made using the following process,

Firstly, the orbit period (Torb) shall be calculated. For RGT orbit it is defined as,

FIGURE 2: SHAPE OF THE DESIGNED ORBIT

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Torb =TEarth

KOrbPerDay [1]

Where TEarth is the Earth rotation period and KOrbPerDay the number of cycles per day. From this expression the altitude of the orbit is obtained.

Torb = 2π√a3

μ [2]

The inclination of the Sunsynchronus orbit is obtained from [3]. This expression makes the perturbance of the gravity field (J2) to change the RAAN of the orbit in way that the angle between the vector normal to the orbit plane and the vector that goes from the sun to the Earth remains constant.

−3

2J2√

μ

a3(

REarth

p)

2

cos(i) =2π

24 × 3600 × 365.25 [3]

Where a is the semi-major axis, μ is the gravitational constant, J2 is the perturbation of the gravity field and p=a(1-𝑒2) where e is the eccentricity of the orbit1.

In order to suit a range of possible altitudes of deployment from the launcher, two limit orbits are defined. This is important in order to know the limit accesses time to the ground stations.

𝐾𝑂𝑟𝑏𝑃𝑒𝑟𝐷𝑎𝑦 𝑇𝑜𝑟𝑏 Altitude Inclinationn

Limit 1 15,5 5574 417 97,0909201

Limit 2 15 5760 567 97,6579919 TABLE 4: LIMIT ORBITS FOR THE MISSION

The design of these limits is based on two criteria:

1. Orbits with an altitude higher than 600 km Will not deorbit within the legal limits.

2. Orbits whit an altitude lower than 300 km have a lifetime lower than the specified in the requirements.

Figure 3 and Figure 4 show the evolution of the height of the orbit during the deorbitation phase. As it can be seen, both orbits are within the permitted time mission upper limits.

1 e=0 because the orbit is circular.

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Accesses times are analyzed in the CS sections, focusing on the Kiruna ground station base.

FIGURE 3: DEORBITATION TIME FOR A CIRCULAR ORBIT OF ALTITUDE 567 KM

FIGURE 4:DEORBITATION TIME FOR A CIRCULAR ORBIT OF ALTITUDE 410 KM

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PAYLOAD (PL)

The payload study objective was to understand and size the data created during the mission and the storage requirement in the satellite. The payload module generates:

• Each minute: 1080p photograph. • Each hour: 10s 1080p video.

Without further information from the client, the following assumptions has been made:

• Photographs and videos taken have a color depth of 8 bits per pixel. • Videos have 30 fps. • When the video is taken, no photograph is made.

Thus, in 1 hour of mission 59 photos and 1 video are made. These assumptions lead to the results of the total data creation per day:

Photographs Video Resolution FULL HD (1080x1920) FULL HD (1080x1920) Pixels 2073600 2073600 Color depth (bit/pix) 8 8 fps N/A 30 File size (bit) 16588800 4976640000 File size (MB) 1.978 593.26 Size required in 1 hour (MB)

709.94

TABLE 5: PL DATA GENERATION

Knowing the periods without communication, it is possible to extrapolate the memory needed to storage all data created in those periods. The maximum time without ground link in the worst-case orbit (for this parameter) is of 11.3 hours, so it is needed a minimum memory capacity of 7.9 GB.

SYSTEM OPERATIONS MODES

The following four system operations modes have been considered in order to minimize the risk along the mission: commissioning, nominal, safe and latency. They differ from each other depending on the needs of the required subsystem for the different scenarios.

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SPACE PROPULSION SUBSYSTEMS (SPS)

During the project a first analysis was made in order to satisfy the legislation regarding the maximum orbital time that a mission could last. The idea of integrating a rocket engine was to change the operational orbit to a disposal orbit when the scientific mission concluded.

However, this solution was rejected for the following reasons:

• Increasing of the space needed (from 6U to 8U) • Increasing of the spacecraft cost. A rocket engine • Increasing the mass of the spacecraft, and consequently the launching

prize.

The alternative solution, and the chosen one, is to disintegrate the satellite during the deorbitation.

ATTITUDE, DETERMINATION AND CONTROL SUBSYSTEM (ADCS)

The first step in the preliminary design of the ADCS is to identify the spacecraft geometry and the operational attitude and orbit. In this case, the spacecraft will be a tetrahedron of 6U Earth oriented in a dawn-dusk sun synchronous orbit.

The ADCS will operate since the deployment of the spacecraft from the launcher and until its deorbitation. During its lifetime, the ADCS shall:

• Detumble the satellite after deployment, pointing the spacecraft in the nominal attitude.

• Provide high precision attitude determination and pointing accuracy during the operation phase (0.5º degree pointing accuracy).

• Redirect the vehicle to control deorbiting.

With these requirements, the control modes of the ADCS can be defined as follows:

COMMISSIONING NOMINAL SAFE LATENCY

Payload

(Pump&Probes)Payload (Pump&Probes&Imaging)

Payload

(Pump&Probes)

Payload

(Pump&Probes)

Data handling & OBC Data handling & OBCData handling &

OBC

Data handling &

OBC

EPS EPS EPS EPS

Attitude control - - Attitude control

- Transmitter&Receiver Receiver Receiver

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CONTROL MODES

1º Deployment

Control and stabilization of possible deployment spin. (Detumbling after deployment)

These modes are used to recover from a external unexpected disturbance or other

emergencies 2º Acquisition Attitude determination and correction before deployment mode

3º Normal Determine and maintain nominal attitude during the mission

4º Fine pointing

Provide fine pointing when required (Payload operation 0.5º degree)

5º Reorientation

Rotate the spacecraft to achieve special targets (avoid sun radiation, augment drag incident surface)

TABLE 6: CONTROL MODES OF THE SATELLITE

Knowing the scenarios where the ADCS will operate, an estimation of possible disturbances and control torques that it will deal with is made.

The disturbance torques that have been considered are:

● Solar Radiation Pressure Torque:

𝑇𝑠𝑝 =𝑆

𝑐𝐴𝑠𝑝(1 + 𝑞)|𝑐𝑠𝑝 − 𝑐𝑔| [4]

𝑆 Solar constant 1367 𝑊/𝑚2 𝑐 Speed of light 3 · 108 𝑚/𝑠

𝑞 Reflectivity coefficient 0.6

𝐴𝑠𝑝 Projected area in Sun direction 0.06 𝑚2

|𝑐𝑠𝑝 − 𝑐𝑔| Distance between the center of solar pressure

and the center of gravity 0.5∗ 𝑚

*Estimated from similar spacecrafts

TABLE 7:SOLAR RADIATION PRESSURE TORQUE COEFFICIENTS

● Aerodynamic Torque:

𝑇𝑎 =1

2𝜌𝑎𝑡𝑚𝐶𝐷𝐴𝐷𝑉𝑝

2|𝑐𝑎 − 𝑐𝑔| [5]

𝜌𝑎𝑡𝑚 Atmospheric density at periapsis ~1 · 10−11 𝑘𝑔/𝑚3

𝐶𝐷 Drag coefficient 2.5

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𝐴𝐷 Drag area 0.03 𝑚2

|𝑐𝑎 − 𝑐𝑔| Distance between the center of gravity and the center of pressure

0.5 𝑚

𝑉𝑝 Spacecraft velocity at the periapsis 7650 𝑚/𝑠

TABLE 8: AERODYNAMIC TORQUE COEFFICIENTS

● Gravity-Gradient Torque:

𝑇𝑔 =3𝜇

𝑟𝑝3

|𝐼𝑚𝑥 − 𝐼𝑚𝑖𝑛| · 𝑠𝑖𝑛 2𝜃 [6]

𝜇 Standard gravitational parameter (Earth) 3.986 · 1014 𝑚3/𝑠2

𝑟𝑝 Periapsis radius 6.83 𝐸 + 06 𝑚

|𝐼𝑚𝑥 − 𝐼𝑚𝑖𝑛| Difference between the maximum and minimum inertia moment

0.08 𝑘𝑔 · 𝑚2

𝜃 Angular deviation between the longitudinal axis and nadir-zenith direction

60 𝑑𝑒𝑔𝑟𝑒𝑒𝑠

TABLE 9: GRAVITY-GRADIENT TORQUE COEFFICIENTS

● Magnetic Torque:

𝑇𝑚 ≈𝐷𝑚𝜆𝑀𝑚

𝑟𝑝3 [7]

𝐷𝑚 Spacecraft residual magnetic dipole 0.1 − 20 𝐴 · 𝑚2

𝜆 Magnetic latitude 1 − 2

𝑀𝑚 Earth magnetic moment 7.96 · 101515 𝑇 · 𝑚3

TABLE 10: MAGNETIC TORQUE COEFFICIENTS

With these parameters, an estimation of the order of magnitude of the disturbance torques can be made:

Gravity-Gradient Torque 2,93E-04 mN·m

Solar Radiation Pressure Torque 2,19E-04 mN·m

Aerodynamic Torque 1,10E-02 mN·m

Magnetic Torque 1,00E+00 mN·m

Total (Worst case) 1,01E+00 mN·m TABLE 11: ESTIMATION OF DISTURBANCE TORQUE

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In the worst case, the highest disturbance torque is the magnetic torque which is about 1 mN·m.

On the other hand, is very frequent that the launchers deploy the spacecrafts with a stabilization spin that could be a problem, due to the mission requirement (high rotation speed (3RPM) longer than 10 seconds shall be prevented). Therefore, the ADCS must ensure that this possible spin is reduced to the operational range in no more than 10 seconds. In order to make an estimation of the torque needed to stabilize the spacecraft, the angular momentum equation can be formulated:

𝐿 = 𝐼𝑧𝜔 ; 𝑑𝐿

𝑑𝑡= 𝑇𝑒𝑥𝑡 [8]

If the inertia moment and the external torque are assumed constant and the external torque and the angular momentum vector are assumed parallel, the evolution of the angular speed with the time will obey the relation:

𝜔(𝑡) = 𝜔0 −𝑇𝑒𝑥𝑡

𝐼𝑧𝑡 [9]

So, the torque needed to achieve a 𝜔𝑚𝑥 = 3 𝑟𝑝𝑚 in a time of 𝑡𝑓 = 10𝑠 with an initial angular speed 𝜔0 is:

𝑇𝑚𝑖𝑛 = (𝜔0 − 𝜔𝑚𝑥)𝐼𝑧

𝑡𝑓 [10]

For a generic spin deployment (Ariane 𝜔𝑚𝑥 = 7.5 𝑟𝑝𝑚 ) the minimum torque needed to stabilize the spacecraft is of 2.83 mN·m.

Having in consideration the high precision pointing requirement and the quick performance needed if there were a perturbation that would induce a greater spin than 3 RPM, a set of reaction wheels in the three axes has been selected as the principal ADCS actuators. The selected set has been the CubeWheel set of CubeSpace (3) which have different sizes available:

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The critical performance parameter in this mission is the maximum torque produced which in this case is determined by the torque needed in the deployment stabilization. As it has been presented, the maximum torque has to be at least about 3 mN·m and none of these sizes achieve it so the team concluded to use a 45º degree configuration of 3 wheels of size S+ in order to combine the torque of them all at the same time. With this configuration the set of 3 wheels will achieve a maximum torque of 4.88 mN·m. The problem of this solution is that if we use the 3 wheels to produce the maximum torque in them all then there will not be an auxiliary wheel to dump moment from the others, so the solution suggested is to use magnetorquers to carry out this task. Although, it is considered in the risk analysis the possibility of carry one redundant wheel.

In the other side, to achieve the pointing accuracy, the ADCS will need a high precision sensor to determine the attitude. The sensor selected that can achieve this precision is a star tracker for fine determination and redundant sensors for normal operation.

With all this in mind, the solution proposed is the 3-Axis integrated ADCS of CubeSpace due to its wide heritage because it has flown in more than 40 missions which one of them is the QB50 cubesats in which the ADCS has a similar requirements as ours (4). Furthermore, this integrated system met all the requirements imposed and its control modes consider the modes needed for our mission.

The configuration selected of the system would be:

• ADCS OBC: CubeComputer with the Control software • CubeControl: Sensor and actuator interface board. • 3 CubeWheel Size S+ mounted in 45º degree assembly • 10 Coarse Sun Sensors

TABLE 12: PERFORMANCE OF THE CUBEWHEEL REACTION WHEELS

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• 3 Magnetorquers • 1 three axis magnetometer • 1 Star Tracker

The company CubeSpace offers a Star Tracker compatible with the ADCS integrated named CubeStar but the only problem with this sensor is that it has a 8 TRL due to it has not flown yet, so the solution proposed is to mount as redundant sensors the CubeSense (Fine Sun and Earth optical sensors) which achieve a 0.6 degree determination precision in the sunlit part of the orbit (all the orbit in our mission).

With this solution, the ADCS can stabilize in time a deployment spin about 10.5 RPM which is a higher spin than the commonly deployment spins for cubesats. Also, the spacecraft will be able to meet the pointing requirements and to operate in all the control modes defined.

COMMAND AND DATA HANDLING (C&DH)

Due to the importance of this subsystem, a very solid solution was needed. The malfunction of this subsystem would imply the complete failure of the mission.

As a first solution, the ADCS OBC was considered as the spacecraft OBC but after contacting with the manufacturer client service (Beniot Chamot from CubeSpace) the solution was changed because they did not think that their OBC could accomplish all the data handling such as control the payload data, comprise it and coordinate all the systems, so the ADCS OBC is only considered to run the ADCS control software due to his low frequency microcontroller. Therefore, the OBC choose was the one of ISIS (5) with a fly heritage since 2012 lowering the risk magnitude of failure and with the hardware requirements needed for the correct data handling.

COMMUNICATION SUBSYSTEM (CS) AND GROUND SEGMENT (GS)

In order to consider the communication subsystem of the spacecraft, the requirements, the payload data-generation, the orbit design and the usable antennas must be studied at the same time. The payload data generation and the requirements are imposed by the client. The orbit design will be established, among others, depending on the performance of the communication subsystem. However, a range of altitudes has been specified by the MA chief. Finally, the usable antennas, both in ground and space will be considered during the whole concurrent design, in order to perform the analysis knowing estimated numbers.

From the requirements and the previous payload analysis, it is known that 550 MB of compressed data must be transmitted per day. Moreover, the

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requirements establish that the ground usable antennas are the ones of the ESA Core Network (see Appendix III: ESA Network); and the range of altitudes proposed by the MA chief are only LEO, leading to analyze only the non-deep space antennas.

Once the data to be transmitted and a range of altitudes are stablished, the analysis can be performed. In order to accept the correct functioning of the subsystem, it will have to perform its mission in the worst case:

• Largest distance to the ground base antenna (more free space losses). • Largest angle to the ground base antenna (worst gain for the antennas). • Lowest fly-by time (higher data rate required).

The range distance proposed by the MA chief is between 410 and 570 km. Both cases will be considered, due to the importance in one of them of the data rate and the distance in the other. If both cases are achievable, it will be considered the correct performance of the subsystem. The whole analysis will be done with

the 410 km altitude case and then extrapolated to the other.

Figure 5 from (2) is considered in order to calculate the angles and distances, when point target is considered the ground antenna. With a typical value ε=15° and

the orbital altitude H=410km:

η (°) λ (°) D (km) 65 9.8 1200

The Band of transmission and reception has been chosen considering the ground stations and the CubeSat communications COTS. Moreover, in order to use just one antenna and one radio, reception and transmission will be done with the same frequency: S-band.

FIGURE 5: ANGLES AND DISTANCE CONSIDERED FOR THE CS ANALYSIS.

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With the frequency band and the η established, the following assumptions are made, considering the state of the art (6) and being conservatives with data:

1. Radio output power will be 1W (30 dbm). 2. The antenna will have a gain of 0db 2. 3. Losses radio-antenna will be 3db. 4. Pointing losses will be 0.5db.

With this assumption, the output power of the antenna is of 26.5 dbm. Considering the distance and the transmitting frequency band, free space losses can be calculated. Moreover, the propagation losses and the polarization are also considered. All these losses are -163 db.

At this point, the ground antenna is receiving the signal, so the ground station must be considered. Studying the possible ground stations (7), the diameter of the ground antenna will be approximately 13 m. With an efficiency of 60%, and band-S correspond to a gain of 46 db. Moreover, the system temperature noise is 135K, corresponding to 21.3 db-K.

Considering a typical radio sensibility of the ground antenna of -110 dbm, summing all the losses described before, knowing the output transmission power of the satellite and the gain of the ground antenna, it is evident that there is enough power margin (~19dbm).

The receiver radio is capable of receive the signal; however, it must be capable of understand it. For this, the carrier to noise power density (C/N0) is calculated. It is important to note that until this point the data-rate has not been considered. Moreover, the only variation between the orbit of 410km and the one of 570 km is because of the free-space losses, where the distance is relevant. Therefore the 570 km orbit C/N0 can be easily calculated without practically varying the data from the previous orbit3.

C/N0 (dBHz) 410 km 570 km

86.9 84.6 TABLE 13: CARRIER TO NOISE POWER DENSITY FOR DIFFERENT ORBITAL ALTITUDES.

From this point, the data rate and hence the fly-by time will be critical. From the shape of the orbit (heliocentric) and the RAAN, it is only considered the Kiruna ground station transmission. This will lower the DOCs and ease the operations. With the two orbital options proposed, the connection times are periodical. The

2 The gain of the antenna has been established considering the worst transmitting angle. 3 The power margin for the 567km orbit is about 17 dBm.

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410 km is 2-day periodical and the 570 km, 1-day periodical. The timeframes with Kiruna are outlined in

Timeframes 410 km 570 km

DAY 1 719 s 1560 s DAY 2 686 s 1560 s

TABLE 14: TIME FRAME WITH KIRUNA GROUND STATIONS FOR DIFFERENT ORBITAL ALTITUDES

With this information, the data rate necessary in each case can be calculated. Leading in the worst case of both orbits to the results,

Data rate 410 km 570 km

6.73 Mb/s 2.96 Mb/s TABLE 15: DATA RATE FOR EACH OF THE ORBITS

With all this information, the Eb/No is calculated for each case.

𝐄𝐛/𝐍𝐨 410 km 567 km 18.6 dB 19.9 dB

TABLE 16: EB/NO FOR EACH OF THE ORBITS

Stablishing a typical BER of 10−5 and a margin of 3dB, only the FSK modulation could be a problem, so other modulation will be chosen to prevent it.

The uplink transmission follows the same design process, not supposing a great variation in the design due to the small size of the data that is needed to transmit to the satellite. This analysis was made for the worst case and is possible to make the link. This means that during the fly over the ground base the data rate can be augmented, and the signal will arrive to ground with more power.

The solution proposed is the state-of-the-art radios (6) from Endurosat. One transmitter band-S radio and another receiver band-S radio (8). Both radios can have up to 32GB of memory, accomplishing the requirement of the data storage for the long periods without ground communication. Moreover, as shown in the datasheets, the output power from the radios are greater than 1, so the data rate of the mission could be augmented.

The antenna chose is a patch one, in order to have a positive gain for a wide-angle spectrum. The S-BAND Patch Antenna Type II from EnduroSat (9) have been chosen due to the variety of possible layouts and its peak gain. This antenna supports a QPSK modulation, so it suits the requirements stablished before.

ELECTRIC POWER SUBSYSTEM (EPS)

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The electrical power subsystem (EPS) provides, stores, distributes and control spacecraft electrical power. In order to size each component of this subsystem we must identify the electrical power loads for mission operations at the beginning-of-life, BOL, and end-of-life, EOL.

For many missions, the EOL power demands must be reduced to compensate the solar array performance degradation. In this case, this difference between BOL and EOL is not important, because of the short life of the mission (three months). The average electrical power needed at EOL determines the size of the power source. Then, a detailed power budget, at different stages of the mission, must be done.

Nevertheless, the peak electrical power budget suits with precision all the scenarios of the mission. This is because the mission last only three months and the designed orbit allows to keep the sunlight direct incidence on the solar panels during the whole mission, avoiding eclipses.

Therefore, the detailed power budget, considering all the subsystems and the requirements of the mission, is the following one:

Subsystem Pmax(W)

Payload (Imaging) 5

Payload (Pump and Probes) 1,5

Payload (Internal heating) 0

Propulsion 0

Attitude Control 0,85

Communications 7,20

Data Handling 0,4

Thermal 0

Power 0,16

Structure 0

Total Consumption 17,38

TABLE 17: ELECTTRICAL POWER BUDGET

Some subsystems performance shall be clarified. The payload subsystem is divided this way because its ON and OFF modes of its subcomponents:

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• Imaging, it is not always working which implies that the the batteries will charge, if required, while this part of the payload does not demand energy. However, the peak power demand will be always available to take pictures or videos, when necessary.

• The pump system and probes part will be always ON. • Internal heating, as it has been said before, will not be turned ON because

eclipses will be avoided during the mission.

Regarding the propulsion subsystem, different ideas were proposed at the beginning of the design in order to end the mission in a disposal orbit. However, it will not be necessary because, as it has been clarified, the satellite will be disintegrated during its reentry into the Earth atmosphere.

Finally, regarding Table 17, an extra margin is added to the total amount of power based on design maturity which is 15%.

Once the power budget is defined, the design procedure to size and choose solar panels and batteries is explained:

The first elements to size are the photovoltaic solar cells, which convert incident solar radiation to electric energy. These components are well-known, reliable and very common in space industry.

Firstly, the mission life and the average power requirement must be defined. In this case, they are respectively 3 months and the peak power requirement from the power budget given in Table 17, 17.38 W.

Once these two key parameters have been established, the total power that the solar array must provide during daylight (Psa) can be determined using [11]:

Psa =

PeTe

Xe+

PdTd

Xd

Td [11]

In this case, the term regarding the eclipse is suppressed, due to the chosen orbit. Therefore, the power that the solar array must provide is [12]:

Psa =Pd

Xd=

17.38

0.85= 20.45 W [12]

Where 𝐏𝐝 is the power requirement and the 𝐗𝐝 is the efficiency of the path directly from the arrays to the loads, which usually takes the value of 0.85 for direct energy transfer. The previous formula shows the total power that must be provided in one orbit to fulfill the spacecraft power requirements.

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Concerning the cell type, multijunction solar cells from Gomspace have been chosen because of the high efficiency required for the design. It is considered that an efficiency of 28% is reached by these cells, although in the datasheet a 30% efficiency is specified (10). Besides this efficiency, a typical coefficient for the inherent degradation 𝐈𝐝 of 0.77 is used in [13].

PBOL = PSUNηIdcosθ [13]

Where PSUN is the solar illumination intensity (1367 W/m) and cos θ is referred as the cosine loss. In this case, the incidence angle 𝛉 between the vector normal to the surface of the array and the Sun line is 0 (because of the orbit). Therefore, the Sun rays will be always perpendicular to the solar array surface and provides maximum power. Then the power at BOL is,

PBOL = 1367 · 0.28 · 0.77 · 1 = 294,73 W [14]

Life degradation Ld factor is the next parameter to be considered. This implies the degradation of the performance of the solar array during the mission. It can be estimated using [15].

Ld = (1 − c)tf = (1 − 0.0005)0.25~1 [15]

Where c is the degradation per year (0.0005 for multijunction cell type (2)) and tf is the satellite lifetime (3 months= 0.25 years). The array performance per unit of area at the end of life is then,

𝑃𝐸𝑂𝐿 = 𝑃𝐵𝑂𝐿𝐿𝑑 = 294.69 𝑊 [16]

Finally, the solar array area, 𝐀𝐬𝐚 required to support the spacecraft power requirements is given by.

Asa =Psa

PEOL= 0,069371m2 = 693,71cm2 [17]

Which are equivalent to 7 solar cells. However, in order to compensate the centre of gravity alteration, 8 solar cells are chosen. This extra solar cell will be also usable in terms of redundancy of the subsystem system.

With regards to the energy storage, several subsystems have been determined as necessary during the launch. The batteries have been dimensioned considering three hours of launching (as maximum time) to have an autonomous

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satellite during this period and to keep the following subsystems active to develop their function:

• Pump and probes from the payload to ensure life (3 hours) • Attitude control to achieve the desired objective after launch (10 seconds) • Data handling to interpret and give corrections to attitude control (10

seconds) • EPS to provide energy (3 hours)

One of the most important parameters which characterize a discharge period is known as the Depth of Discharge (DoD), which will be 25%. Once the average DoD is established, determining the different currents that each load needs, the total capacity of the batteries (𝐂𝐫) can be done by using the following expression.

Cr =Pe · Te

DoD · N · Xe= 15,33 Wh = 1,92Ah [18]

Where N is the number of non-redundant batteries (2 in this case) and Xe the efficiency of the paths from the battery to the individual loads, which is 65%. To obtain the battery capacity in ampere-hour, the previous result must be divided by the bus voltage (8 V).

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Batteries can be connected in series to increase the voltage or in parallel to increase the current. In this case, the solar cells are connected in series two by two in order to reach 8V in total, and in parallel also in pairs, in order to transfer the maximum current (2 A) to the batteries connected the following way:

MECHANICAL DESIGN AND STRUCTURE (MD&S)

STRUCTURE

The Structure shall be compatible with the mission, and not bigger than 8U. In this project, it has been selected a 6U dimension and the selected structure if from the ISIS developer, based on the CubeSat standards. It has a flight heritage since 2015, therefore, it is in an advanced grade of development. It is mounted in several missions, such as the Brik II of the Royal Netherlands Air Force, not yet launched, and both Hiber-1 and Hiber-2, both in orbit.

This structure allows the mounting of several PCBs, in different directions. For simplicity, in this phase of the predesign, the secondary structure (in which will

FIGURE 6: DIAGRAM OF THE ELECTRICAL POWER CIRCUIT

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be mounted the different PCBs) will not be displayed, just the PCBs and the components. Finally, the detachable shear panels allow the proper access to the mounting of the different components.

CONFIGURATION

For the configuration department, its objective is to arrange all the needs from the different departments and be able to fit the different elements in the desired structure. Therefore, the following restrictions have been considered:

1. Payload: it will be formed by 3U forming a L shape. For the configuration details, it is just a L-shaped box

2. Attitude, Control and Determination Subsystem (ACDS): 2.1. 4 Inertia Wheels, mounted at 45° and the further as possible from the

centre of the satellite 2.2. 2 magnetorquers 2.3. 2 sensors for the satellite orientation determination, one facing the sun and

the other facing the Earth. Therefore, the front solar panel (which will be facing the Sun) must have a hole in it to allow the sensor to see the sun. The same problem is considered in the bottom shear panel for the other sensor

2.4. 1 Star Sensor. It will give the precision imposed in the requirements. It must face the stars and must not see the sunlight, nor directly neither by reflection. It is mounted then in the back face of the satellite

FIGURE 7:6U STRUCTURE

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3. Command & Data Handling: 3.1. 1 S-Band Antenna: facing the Earth 3.2. 1 Receiver + 1 Transmitter: both elements may be mounted together in

order to get a correct performance of the transmission system 4. Electric Power Subsystem:

4.1. 6U faces of solar panels. These will always face the Sun 4.2. 2U faces of additional solar panels. These will be deployed once in orbit 4.3. 1 P31 Battery + 1 redundant P31 battery

5. General: 5.1. On-Board Computer

FIGURE 8: DETAILED VIEW OF THE COMPONENTS

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For this project, the configuration has been developed both in CATIA and Idm-CIC. In CATIA, the model has been obtained with the components from the manufacturers themselves, so that the design is more realistic. In some cases, it has been necessary to contact the companies to obtain the Step files (Endurosat). On the other hand, for the Idm-CIC, the model has been developed trying to simulate as real as possible the components or obtaining them from the own Catalogue of the Idm-CIC, changing it in order to adapt them to the selected components.

DEPLOYMENT OF SOLAR CELLS

The 2U top solar panel mounts a mechanism based on springs that enables the deployment at 90º for the solar array.

FIGURE 9: DEPLOYMENT MECHANISM

This mechanism gives an accuracy less than 0.3º and a range of up to 180º. It has been used in the ESA’s Solar Orbiter mission and it currently has TRL 8.

This asymmetrical deployment of the solar panel leads to a momentum in the satellite, which will be corrected by the ADCS.

MASS BUDGET

For the definition of the masses, as it is used COTS, all the components previously mentioned are obtained from several companies specially dedicated to Space products. Therefore, it is possible to get, as a first approximation for the predesign, the mass of each component and then the whole mass of the satellite. Depending of the development of the products, a mass margin for each component shall be added:

• Fully developed: 5% • To be modified: 15% • To be developed: 20% • As mentioned, most components are fully developed and have flown in

several missions, therefore the margin for most of them will be low. In the

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following table, a complete summary of all the masses and its margins has been arranged. For more information of all components, it can be found in the Idm-CIC file of this project.

Quantity Mass [kg] Mass with margins [kg]

ACDS Inertia Wheels 4 0,06 0,063

Torquers 2 0,036 0,038

Sensors 1 0,144 0,151

Cube Sense 1 0,16 0,168

Cube Star 1 0,055 0,058

Payload 1 7 7,35

Structure Principal 1 1,1 1,155

Solar Panels 2U 1 0,1 0,105

3U 2 0,15 0,34

6U 1 0,3 0,315

Batteries 2 0,2 0,21

OBC 1 0,077 0,081

Antena 1 0,15 0,158

Transmitter 1 0,08 0,084

Reveiver 1 0,08 0,084

TOTAL 10,1 10,9

TABLE 18: MASS BUDGET

With the masses defined, it is possible to get the centre of gravity of the satellite and its Inertia moments, which will be needed for the ADCS subsystem.

As shown inFigure 10, the Centre for the whole system is very near to the centre of the payload. This is because the payload is very heavy, compared to the rest subsystems (7kg out of 9kg for the complete satellite).

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THERMAL CONTROL SUBSYSTEM (TCS)

The preliminary design process of the thermal control subsystem has been divided in two steps. Firstly, the identification and characterization of the parameters that are going to play an important role in the thermal model, which are the followings:

1. Radiation sources, established as the solar radiation, the earth albedo and the earth irradiance. All of them have been characterized by the radiation fluxes defined in the ECSS (11).

Solar radiation flux (Solar constant 𝑆)

1319.7 – 1412.5 𝑊/𝑚2

Earth albedo (𝑎) 5% - 60% Percentage of sun radiation reflected

Earth irradiance (𝐼𝑅) 150 - 350 𝑊/𝑚2

TABLE 19: RADIATION SOURCES

FIGURE 10: CENTRE OF GRAVITY AND INERTIA OF THE SATELLITE

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Due to the variation of the fluxes, two cases are considered: the hot case, with the hottest values and the cold case, with the coldest ones.

2. Internal dissipation from the components of the spacecraft and the operational temperature range of the components and the payload:

Dissipation (W) Temperature range (ºC)

ADCS 0,7 -10 to 70

OBC 0,3 -25 to 65

BATTERIES x2 0,3 -40 to 85

RADIO 2,8 -40 to 85

PAYLOAD 0,7 11 to 15

TOTAL 4,8 -10 to 65 TABLE 20:DISSIPATION OF INTERNAL COMPONENTS

The payload has its own thermal control system and can operate in the normal temperature range of the other components, so the operational temperature range considered in the TCS design has been the most restrictive for the components. In the other hand, the dissipation considered has been in the normal operation of the systems.

3. The simplified orientation and the geometry of the spacecraft in the nominal orbit: the spacecraft is modeled as a 6U tetrahedron in the down-dusk Sun-synchronous orbit. One of the 6U sides is oriented towards the Sun and a 2U face will be always oriented to the Earth surface. This orientation determines that the angle of incidence (beta angle) of the solar radiation is going to be ~0º during the mission lifetime.

FIGURE 11: SCHEME OF PRINCIPAL AXES AND RADIATION ORIGIN

Once the identification and characterization of these parameters are made, it has proceeded with a simplified thermal model thought for a preliminary estimation

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of the equilibrium temperature that will be in the spacecraft under the previous conditions. The mathematical model used to compute the estimations states the spacecraft as a tetrahedron of 22 faces of 10cm x 10cm and the vehicle interchanges radiation with the external environment through them until an equilibrium temperature is reached. In the equilibrium the heat emitted and received are equals:

𝑄𝑖𝑛 = 𝑄𝑜𝑢𝑡 [19]

And due to the different sources of radiation the heat received has the next components:

𝑄𝑖𝑛,𝑠𝑢𝑛 = 𝐴𝑆𝛼𝑆𝑆 𝑄𝑖𝑛,𝑎𝑙𝑏𝑒𝑑𝑜 = 𝐹𝑂𝑉 · 𝐴𝑒𝛼𝑒𝑎𝑆

𝑄𝑖𝑛,𝐼𝑅 = 𝐴𝑒𝜖𝑒𝐼𝑅 𝑄𝑖𝑛,𝑑𝑖𝑠𝑖𝑝𝑎𝑡𝑒𝑑

[20]

Where

• 𝑄𝑖𝑛,𝑠𝑢𝑛 is the radiation received by the sun, where 𝐴𝑝 is the area of the faces to sun, 𝛼𝑝 is the absorbance of the material of the faces towards the sun and S is the solar constant.

• 𝑄𝑖𝑛,𝑎𝑙𝑏𝑒𝑑𝑜is the radiation absorbed from the Earth albedo, where 𝐴𝑒is the area of the 2 faces pointing Earth and 𝛼𝑒 the absorbance of the material of these 2 faces, and FOV is the factor of field of view (1.5 for an altitude of 450 km)

• 𝑄𝑖𝑛,𝐼𝑅 is the infrared radiation absorbed from the Earth by the faces towards Earth and 𝜖𝑒is the emissivity of the material of these faces.

• 𝑄𝑖𝑛,𝑑𝑖𝑠𝑖𝑝𝑎𝑡𝑒𝑑 is the dissipation of energy in form of heat of the internal systems.

In the other hand, the heat emitted in form of infrared radiation is modeled by the Stefan Boltzmann law:

𝑄𝑜𝑢𝑡 = 𝜎𝑇𝑒𝑞4 ∑ 𝜖𝑖𝐴𝑖

22

𝑖=1

[21]

Where 𝜎 is the Stefan-Boltzmann constant and the summation is for all the external faces. From the last equation the equilibrium temperature can be solve.

In this case, we have the next possible materials to implement in our spacecraft:

𝛼 𝜖

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Solar Arrays 0,93 0,81

Coating (MLI) 0,30 0,6

Aluminum sheet metal panels 0,55 0,85

TABLE 21: POSSIBLE MATERIALS FOR COATING THE SPACECRAFT

After a several iterations of different possible configuration, the configuration selected has been the 6 faces towards the sun covered with solar arrays and the other 6 faces in the opposite of the Sun with multilayer insulation protection (MLI). The rest of the lateral faces are going to be covered with redundant solar arrays in spite of the temperature is a little higher but with these panels more power could be generated just in case. With this configuration, the heats absorbed, and the equilibrium temperature of the spacecraft are:

Hot case Cold case 𝑄𝑖𝑛,𝑠𝑢𝑛 (𝑊) 78.82 73.64

𝑄𝑖𝑛,𝑎𝑙𝑏𝑒𝑑𝑜 (𝑊) 13.94 1.09 𝑄𝑖𝑛,𝐼𝑅 (𝑊) 8.90 3.81

𝑇𝑒𝑞 (º𝐶) 34.35 11.75 TABLE 22: TEMPERATURE DEFINITION FOR THE HOT AND COLD CASES

This analysis has been done to the entire spacecraft, but due to the payload has its own thermal control system, another thermal model has been studied to estimate the temperature that will be in the 3U outside of the payload where the instruments will be operating. With this goal, the compartment has been modeled as a 3U L shape cube with the correspondent materials in the external faces.

FIGURE 12: DEFINITION OF THE INTERNAL ARRANGEMENT

In this case the radiation absorbed from the Sun will be the half because 𝐴𝑆 will be only 3 faces instead of 6. In the other hand, the albedo radiation and the Earth infrared radiation absorber will be the same. Also, there will be another term in the radiation absorbed which will be the infrared radiation emitted by the payload at 13 degrees which can be calculated by the Stefan-Boltzmann law. In

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this scenario the heat absorbed and the equilibrium temperature of the 3U compartment will be:

Hot case Cold case 𝑄𝑖𝑛,𝑡𝑜𝑡 (𝑊) 67.05 43.16

𝑄𝑖𝑛 𝑓𝑟𝑜𝑚 𝑝𝑙𝑑 (𝑊) 11.48 10.90 𝑇𝑒𝑞 (º𝐶) 36.59 11.19

TABLE 23: DEFINITION OF THE HOT AND COLD CASE

With this configuration the temperature operation of the components is inside the operational ranges and there is no need of carry any active thermal control in the spacecraft such as heaters or coolers but as it is discussed in the risk assessment later there will be a redundant heater to operate in case of any eclipse occur.

Finally, to assure the temperature inside the payload box is in the limit range it has been analyzed by a single component of Al shielding receiving only infrared radiation from the interior of the spacecraft:

FIGURE 13: SCHEME OF THE RADIATION ABSORTION OF THE PAYLOAD MODULE

To compute the irradiance in the spacecraft interior the Stefan-Boltzmann law can be formulated assuming the spacecraft at the equilibrium temperature for the previous configuration and with an average emissivity from its surfaces:

𝐼𝑅𝑠𝑝 = 𝜖𝑎𝑣𝑔𝜎𝑇𝑒𝑞4 [22]

With this analysis the temperature obtained inside the payload box in the two cases considered is:

Hot case Cold case 𝑄𝑖𝑛,𝑡𝑜𝑡 (𝑊) 53.31 50.57*

𝑇𝑒𝑞 (º𝐶) 14.46 10.42 TABLE 24: HOT AND COLD CASE FOR THE PAYLOAD MODULE

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*Considering the internal heaters in the payload

At the view of this analysis results, we can conclude that the equilibrium temperature inside the payload with this configuration meets the operational range.

LAUNCHER

The launcher selected for this mission is the Rocket Lab Electron rocket (12). The design of Electron is optimized to reach sunsynchronous orbits with multiple small payloads. Figure 14 from the Electron User’s Guide relate de altitude with the payload mass for sunsynchronous orbit.

The main reason to choose this launcher is that characteristic of the typical Electron’s payload are very similar to Nanostar. Deviations in the orbit and

deployment margins for this vehicle are:

FIGURE 14: PAYLOAD CHARGE TO SUNSYNCHRONOUS ORBIT FOR ROCKET LAB ELECTRON LAUNCHER

FIGURE 15: DEPLOYMENT MARGINS FOR THE LAUNCHER

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RISK ANALYSIS AND MITIGATION

RISK POLICY

Considering the scientific requirements of the project and the documents provided by the NanoStar consortium two principal objectives are raised:

1. Correct operation of the scientific payload, allowing the correct collection of the data and livelihood conditions for Roscoff worms.

2. Transmission of the scientific data to the ground segment.

In order to well understand the successfulness criteria for the mission, the following definitions must be established:

• Consistent data: It is the trustfully and useful scientific data. It is considered that not consistent data is transmitted when it is corrupted, not coherent or can technically be considered that the data is not valuable. E.g. the payload camera is shooting white pictures for a malfunction. The information from this camera would be considered not consistent.

• Transmitted data: It is the data, consistent or not, that is transmitted to the ground segment.

• Expected data: It is the amount of data expected to be generated for a stablished time period according to the payload specifications in (13).

Therefore, the following successfulness mission criteria are raised:

• Mission success: The mission will be considered successfully if 3 months expected consistent data is transmitted to ground facilities.

• Mission partial success: Mission will be considered a partial success if 2-weeks to 3-months expected consistent data is transmitted to ground facilities.

• Mission failure: The mission will be considered a failure if less than 2-weeks expected consistent data is transmitted to ground facilities.

RISK STRATEGY

During the project, when a SC identifies a risk the following protocol cycle will be set in:

1. The risk will be raised to the SE and the rest of SCs. 2. The SE and the SC will assess the risk accordingly to the tables defined

further on this document, evaluating both the severity and the likelihood. If an agreement is not reached, the consideration of the SE will be the accepted one.

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3. If needed, the SE and the SCs will propose different options for its mitigation, and the SE will choice the optimal one after studying all of them and if the proposed mitigation is enough for acceptance, submit for acceptance.

Because of the nature of the project, there is no customer, so once the team reach the task 9 of the risk management process considered in (14) is reached, it would be considered as an accepted risk.

RISK SEVERITY SCORING

The risk severity is established considering its impact on the performance of the mission.

Score Severity Consequence 5 Maximum Major impact, the system is not capable of performing

none of the technical requirements and the accomplishment of the principal mission objectives is not reached.

4 High Major impact, the system is not capable of performing more than 2 technical requirements and the accomplishment of the principal mission objectives is compromised.

3 Medium Minor impact, the system is not capable of performing more than 2 technical requirements but will accomplish the principal mission objectives.

2 Low Minor impact, the system is not capable of performing 1 technical requirement but will accomplish the principal mission objectives.

1 Minimum Negligible impact, the system can perform all mission requirements and will accomplish all mission objectives.

TABLE 25: SEVERITY SCALE OF RISKS

RISK LIKELIHOOD SCORING

The risk likelihood is established considering the times the failure scenario may occur.

Score Likelihood Likelihood of scenario occurrence E Maximum Certain to occur, will occur one or more times per

project. D High Will occur frequently, about 1 in 10 projects. C Medium Will occur sometimes, about 1 in 100 projects. B Low Will seldom occur, about 1 in 1000 projects. A Minimum Will almost never occur, 1 of 10000 or more projects.

TABLE 26: LIKELIHOOD OCURRENCE OF RISKS

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RISK INDEX AND MAGNITUDE SCHEME

Based on the scoring of the likelihood and the severity explained above the indexing of the risk scenario is shown in Table 27. This indexing has been made in coordination with the different SC in order to reach a common criterion for the interpretation and indexing of the risks identified.

Likelihood E Low Medium High Very High Very High D Low Low Medium High Very High C Very Low Low Low Medium High B Very Low Very Low Low Low Medium A Very Low Very Low Very Low Very Low Low

1 2 3 4 5 Severity TABLE 27: RISK INDEX SCHEME

As is shown, each of the risk index is denoted by one of the following risk magnitudes: Very Low, Low, Medium; High or Very High depending on the risk scenario.

ACCEPTANCE CRITERIA AND ACTIONS PROPOSED

The criteria for acceptance have been very restrictive, only accepting Very Low and Low risks. This is because of the academic nature of the project, not having a big economical budget and focusing on COTS and high TRL components.

The acceptance and the action to be taken for each of the risk magnitudes are summarized in Table 28.

Index Risk Magnitude

Acceptance Proposed action

E4, E5, D5

Very High No The requirements shall be adjusted to avoid the risk.

E3, D4, C5

High No The requirements shall be adjusted to avoid the risk.

E2, D3, C4, B5

Medium No Mitigate the criticality of the risk by eliminating or reducing the likelihood or severity.

E1, D1, D2, C2, C3, B3, B4, A5

Low Yes Control and monitor the evolution of the risk. A mitigation of the risk might be proposed if a major performance improvement is achieved.

C1, B1, A1, A2, B2, A3, A4

Very Low Yes Control and monitor the evolution of the risk.

TABLE 28: INDIVIDUAL RISK ACTION CRITERIA

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RISK ASSESMENT

To proceed with the risk assessment, a method of assessment and a principle of identification had been defined as follows:

METHOD OF ASSESMENT

The risk that are in consideration have been identified by the correspondent subsystem chief of every subsystem and had been communicated to the system engineer. The method that had been used in this phase of the project has been the personal research of each subsystems and their possible failures.

PRINCIPLE

The identification of the risks in each subsystem have been responsibility of each SC. The identification was made based on the experience of each SC and the comparation with similar missions (university CubeSat launches).

The assessment of the severity score of each risk has been made based on the Bernoulli Risk Management Document method. The assessment of the likelihood score of each risk has been made based on the TRL of each component in the case of subsystems; and based on the launch’s successes of the launcher chosen.

The assessment of the likelihood score of each risk has been made based on the TRL of each component in the case of subsystems and based on the launcher successes in the case of the launch risk scenarios.

IDENTIFIED RISK SCENARIOS AND ASSESMENT

The risk register for all the next scenarios has been completed and added to the

1. Solar array deployment failure scenario: this risk scenario had been considered because the possibility of the solar array deployment failure the 2U face resulting in the reduction of the power. Due to this, the severity has been ranked Minimum because the system could be capable of accomplish all the requirements. In the other hand, the likelihood had been ranked to medium despite of the high TRL and their previous heritage of this mechanism due to this kind of mechanisms usually give problems, concluding in a risk index of C1.

2. Solar array malfunction scenario: the consideration of this risk scenario is due to the possibility of the malfunction or the total failure in one of the solar arrays equipped in the spacecraft causing the lack of power and the malfunction of the payload and other´s subsystems. Therefore, we have assigned the maximum severity to this scenario but in the other hand the

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considered solar arrays has a high TRL and the failure of one of them have a low probability, concluding in a risk index of B5.

3. Payload malfunction scenario: this risk scenario had been considered because of the possibility of the malfunction of the payload causing the generation of no consistent data preventing from the accomplishment of the principal mission objective. Because of that, we have assigned a maximum severity to this scenario. By the other hand, the likelihood of this scenario has not been assessed with valid criteria and it will be ranked with further information.

4. Launching system failure scenario: we have considered this scenario due to the possibility of the total failure in the launching system causing the lost of the complete spacecraft and the mission will not be able to be developed, then the severity of this scenario is the maximum severity. Furthermore, the launch vehicle selected, which is the Electron rocket from RocketLab, has done 10 launch with success in 9 of them but the first failure was in the first test flight and was a partial failure without loss of the payload then the probability of success is upper of 90% in this launcher but the probability of the payload loss would be minimum, then the likelihood of this scenario had been ranked to minimum.

5. Nominal orbit parameters not achieved: not achieve the nominal orbit parameters had been considered as a separately risk scenario due to there is not loss of the spacecraft and the mission could been developed but the new orbit could have eclipsed periods and cause thermal anomalies above the expected by the thermal control subsystem preventing from the operation in the design temperature range of the rest of the subsystems and causing the lack of power generated by the solar arrays preventing from to operate correctly the payload and the other subsystems, then the severity of this scenario is ranked to maximum severity and due to the probability of total success of more than 90% the likelihood is ranked to medium.

6. Orbital altitude insertion failure scenario: this scenario is considered because of the possibility of the insertion in a different altitude due to a partial failure in the launch system causing a reduction in the mission duration. This could prevent the compliance of the requirement of minimum life if the altitude is too lower or prevent the compliance of the maximum legal time in orbit if the altitude is too high, and due to this the

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severity has ranked to medium. By the other hand, the likelihood of occurrence of this scenario is the same of the previous scenario and it is ranked to medium too.

7. Onboard computer malfunction scenario: this scenario has been considered due to the possibility of the malfunction of the OBC causing the malfunction of other subsystems like the COM&Data subsystem, the ADCS or the payload operation, then the principal mission objective could not be accomplished. Because of this the severity has been ranked to maximum, but the OBC selected has a high TRL and the likelihood of occurrence of this scenario is minimum.

8. Communication subsystem failure scenario: this individual scenario is considered because of the possibility of the malfunction or total failure of the communication subsystem causing the reduction or the impossibility of the transmission of scientific data and the reception of command data. Because of in this scenario the principal mission objective is compromised the severity has been ranked too high, but due to the heritage and the high TRL of the subsystem selected the likelihood of this scenario is ranked to minimum.

9. Magnetorquer failure scenario: the possibility of the malfunction or total failure of one magnetorquer has made us to consider this scenario, in which the consequences will be the impossibility of dump the momentum build-up on the reaction wheel in the axis of the magnetorquer axis of action. This will cause the limitation of the uses of the reaction wheel, reducing the mission life with control in this axis which compromise the primary mission objective, then the severity has been ranked high. By the other hand, the possibility of failure of these systems has been ranked to low due to the wide heritage of them and of the one selected.

10. Star tracker failure scenario: this risk scenario has been considered due to the possibility of the malfunction or total failure of the star tracker used to achieve the high precision pointing accuracy, then the consequence of this scenario is the non-compliance of this particular requirement lowering the time of accesses about 1% and causing worse operation of the payload (the pointing accuracy will be 1 deg instead of 0.5 deg), then the principal objective will be achieve but one requirement will not be met and due to this the severity has been ranked low. By the other hand, the star tracker selected has no heritage but a high TRL (8), so the likelihood had been ranked high.

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Before the risk scenario identification and assessment, it can be concluded that the overall project risk depends on the correct behavior of the Payload, which is the most critical component and one that has a high severity and an uncertain likelihood. Although, the risks that have a link between them must have a special monitoring mark in order to notify all the SCs if it is a critical risk scenario in the next phases of the project.

RISK TREND

After the assessment and the mitigation of all the risk scenarios identified we have made a representation of the possible evolution of the scenarios between the phase 0 and the phase A.

The payload malfunction scenario has not been considered in this analysis due to the lack of information about its possible failures and mitigations.

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CONCLUSIONS AND FUTURE WORK

As the technical reports of the different subsystems have shown, the Bernoulli team has reached a solution that can perform the mission and the requirements demanded by the client.

Moreover, making an estimation of the costs using a bottom-up method, it is visible that the COTS components reduce the price of the satellite, making it possible for a little company or a university to be made (see. Appendix V: Bottom-Up cost estimation).

The orbit is chosen to prevent the incorporation of heaters and minimize the batteries size. This is one of the strongest points of this satellite, minimizing the electric and mass budget of the spacecraft.

On the other hand, it is shown that communication is possible for all the altitudes ranges. In some of them, the capacity of data that is possible to transmit make possible to augment the quality of the photos and video from the payload, and even augment their recording frequency. This would be traduced in an improvement of the scientific data.

The next step in the project, would be start the manufacturing of the satellite, having to contact all the providers. Moreover, the payload should be well defined, in order to know its failure likelihood and possible mitigation actions.

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APPENDICES

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APPENDIX I: GLOSSARY OF ACRONYMS

Acronym Meaning

ADCS Attitude, Determination and Control

Subsystem C&DH Command and Data Handling COTS Commercial Off-The-Shelf

CS Communication Subsystem CSR Corporate Social Responsibility DOC Direct Operational Cost EPS Electric Power Subsystem fps Frames Per Second GS Ground Segment

LEO Low Earth Orbit MD&S Mechanical Design and Structure N/A Not Applicable OBC On-Board Computer PL Payload PM Project Manager

REACH Registration, Evaluation,

Authorization and Restriction of Chemicals

ROHS Restriction of Hazardous Substances SC Subsystem Chief SE System Engineer

SPS Space Propulsion Subsystem TCS Thermal Control Subsystem TRL Technology Readiness Level WBS Work Breakdown Structure

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APPENDIX II: GANTT DIAGRAM

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APPENDIX III: ESA NETWORK

FIGURE 16: GLOBAL ESA NETWORK. (SOURCE (5))

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APPENDIX IV: RISK REGISTERS

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APPENDIX V: BOTTOM-UP COST ESTIMATION

Considering the Breakdown structure of the mission (Figure 17) and the costs of the different components used, a preliminary cost estimation can be made (

In the communication subsystem, the transceiver was the first design idea, however due to the performance of the endurosat radios and its flying heritage, two radios were chosen. The price of the radios are of 10K$.

FIGURE 17: BREAKDOWN STRUCTURE OF THE MISSION

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FIGURE 18: BOTTOM-UP COST ESTIMATION

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