CDF Study Report: CDF-84(A) February 2009 Page 1 of 267

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 1 of 267

CDF Study Report

Development of a Highly-Capable, Modular Miniaturised Platform

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FRONT COVER

Exploded view of a Nanosat.

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STUDY TEAM This study was performed in the ESTEC Concurrent Design Facility (CDF) by the following interdisciplinary team:

TEAM LEADER S. Santandrea, TEC-SY AOCS J. Grzymisch, TEC-ECC POWER A. De Luca, TEC-EPS S. Wu, TEC-ECN COMMUNICATIONS O. Alvarez, TEC-ETC PROGRAMMATICS A. Frutos Pastor, TEC- /AIV TC M. Braghin, TEC-TCC O. Brunner, TEC-TCC CONFIGURATION R. Klotz, TEC-MSS PROPULSION C. Edwards, TEC-MPE V. Lowe, TEC-MPA COST E. Lamboglia, TEC- RISK A. Harrison, TEC-QQD SYC DATA HANDLING G. Furano, TEC-EDD STRUCTURES A. Shannon, TEC-MSS GS&OPS David Evans, OPS- SYSTEMS K. Gantois, TEC-SY HAS R. Findlay, TEC-SYE N. Murdoch, DG-PI MECHANISMS M. Falkner, TEC- THERMAL G. Chirulli, TEC-MCT MSM MISSION P. De Pascale, OPS- ANALYSIS GFA

Under the responsibility of: S. Airey, TEC-ECC Study Manager J. Kohler, TEC-MMA Study Manager with the technical support of: L. Gerlach, TEC-EPG Solar Generators L. Marchand, TEC-QCT Components and Materials F. Filhol, TEC-QCT Components and Materials M. Sabbadini, TEC-EEA Antennas

The editing and compilation of this report has been provided by: A. Pickering, TEC-SYE Technical Author

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Further information and/or additional copies of the report can be requested from: S. Airey ESA/ESTEC/TEC-ECC Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5655295 Fax: +31-(0)71-5653711 [email protected]

For further information on the Concurrent Design Facility please contact: M. Bandecchi ESA/ESTEC/TEC-SYE Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5653701 Fax: +31-(0)71-5656024 [email protected]

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TABLE OF CONTENTS

1 INTRODUCTION...... 13 1.1 Background...... 13 1.2 Scope ...... 13 1.3 Document Structure...... 14 2 EXECUTIVE SUMMARY ...... 15 2.1 Study Flow...... 15 2.2 Study Objectives...... 15 2.3 Requirements and Design Drivers...... 15 2.3.1 System Requirements ...... 15 2.3.2 Design Drivers...... 16 2.4 Study Approach...... 16 2.5 Technical Conclusions...... 17 2.5.1 Baseline Configuration...... 17 2.5.2 List of Modules...... 19 2.6 Test Mission Scenarios...... 22 2.6.1 Scenario 1 – LEO SSO (w/o Propulsion) ...... 22 2.6.2 Scenario 2 – LEO SSO (w/ Propulsion) ...... 23 2.6.3 Scenario 3 – GTO ...... 24 3 STUDY OBJECTIVES AND APPROACH ...... 27 3.1 Miniature Satellite Background...... 27 3.2 Modularity – The Cubesat Example...... 28 3.3 Study Objectives & Trade-Offs...... 29 3.4 Study Starting Assumptions and Generic Platform Requirements...... 29 4 MISSION ANALYSIS ...... 31 4.1 Requirements and Design Drivers...... 31 4.1.1 Mission Objective and Reference Mission Scenario...... 31 4.1.2 Launch Year and Mission Duration...... 31 4.1.3 Orbit Maintenance...... 31 4.1.4 End of Life Disposal...... 31 4.2 Assumptions and Trade-Offs...... 31 4.2.1 S/C Size and Mass ...... 31 4.2.2 LEO Option: Altitude and Maintenance Cost ...... 32 4.2.3 GTO ...... 33 4.2.4 De-orbiting...... 34 4.3 Baseline Design...... 35 4.3.1 Scenario 1: SSO Orbit Without Propulsion...... 35 4.3.2 Scenario 2: SSO Orbit With Propulsion...... 35 4.3.3 Scenario 3: GTO...... 36 4.4 Budgets ...... 37

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5 SYSTEM ...... 39 5.1 Purpose and Scope...... 39 5.2 System Requirements ...... 39 5.3 System Design Drivers ...... 40 5.4 System Development Activities ...... 40 5.4.1 Requirements and Design Drivers ...... 40 5.4.2 Activity/ Module List...... 41 5.5 Mission Baselines - Scenario 1 (LEO w/o Propulsion)...... 41 5.5.1 System Baseline Design...... 41 5.6 Mission Baselines - Scenario 2 (LEO w/ Propulsion)...... 44 5.6.1 System Baseline Design...... 45 5.7 Mission Baselines - Scenario 3 (GTO)...... 48 5.7.1 System Baseline Design...... 49 6 CONFIGURATION...... 55 6.1 Requirements and Design Drivers...... 55 6.2 Assumptions and Trade-Offs...... 55 6.3 Baseline Design...... 55 6.4 Scenario Study Case 1 - LEO (w/o Propulsion) ...... 57 6.4.1 Design Drivers...... 57 6.4.2 Description...... 57 6.4.3 Modules Selection (With Justification) ...... 59 6.4.4 Achievable Performances/Realised Functionalities...... 59 6.5 Scenario Study Case 2 - LEO (w/ Propulsion) ...... 59 6.5.1 Design Drivers...... 59 6.5.2 Description...... 59 6.5.3 Modules Selection (With Justification) ...... 61 6.5.4 Achievable Performances/Realised Functionalities...... 61 6.6 Scenario Study Case GTO...... 61 6.6.1 Design Drivers...... 61 6.6.2 Description...... 62 6.6.3 Modules Selection (With Justification) ...... 64 6.6.4 Achievable Performances/Realised Functionalities...... 64 6.7 Overall Dimensions ...... 64 6.8 Options...... 67 6.9 Technology Requirements...... 67 7 STRUCTURES...... 69 7.1 Requirements and Design Drivers...... 69 7.2 Sub-System Description...... 69 7.2.1 Interfaces...... 70 7.2.2 Half Wall Panels...... 71 7.2.3 Top Panel...... 72 7.2.4 Bottom Panel...... 73 7.2.5 Panel Brackets...... 74

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7.3 Module Descriptions...... 76 7.3.1 Module 1 (Wall Panel) ...... 76 7.3.2 Module 2 (Half Wall Panel) ...... 77 7.3.3 Module 3 (Top Panel)...... 77 7.3.4 Module 4 (Bottom Panel) ...... 78 7.3.5 Module 5 (Panel Interfaces)...... 78 7.3.6 Module 6 (Harness)...... 79 7.4 Scenario Study Case 1 (LEO w/o Propulsion) ...... 79 7.4.1 Design Drivers...... 79 7.4.2 Module Selection (Justification)...... 79 7.4.3 Mass Budget...... 80 7.5 Scenario Study Case 2 (LEO w/ Propulsion) ...... 80 7.6 Scenario Study Case 3 (GTO Scenario) ...... 80 7.6.1 Design Drivers...... 80 7.6.2 Module Selection (Justification)...... 80 7.6.3 Mass/Power Budget...... 80 8 MECHANISMS...... 81 8.1 Requirements and Design Drivers...... 81 8.2 Sub-System Description...... 81 8.2.1 Interfaces ...... 81 8.3 Module Descriptions...... 82 8.3.1 Module 1: Solar Array Deployment Mechanism (SDM) ...... 82 8.3.2 Module 2: De-orbit Deployment Mechanism (DDM)...... 84 8.3.3 Module 3: Hold-Down and Release Mechanism (HDRM) ...... 86 8.3.4 Module 4: Nano-Terminator De-Orbit Mechanism (NTDM) ...... 90 8.4 Scenario Study Case 1 (LEO Without Propulsion) ...... 94 8.4.1 Design Drivers...... 94 8.4.2 Module Selection (Justification)...... 94 8.4.3 Mass Budget...... 94 8.4.4 Capabilities Provided to a Payload...... 94 8.5 Scenario Study Case 2 (LEO With Propulsion) ...... 95 8.5.1 Design Drivers...... 95 8.5.2 Module Selection (Justification)...... 95 8.5.3 Mass Budget...... 95 8.5.4 Capabilities Provided to a Payload...... 95 8.6 Scenario Study Case 3 (GTO) ...... 95 8.6.1 Design Drivers...... 95 8.6.2 Module Selection (Justification)...... 95 8.6.3 Mass Budget...... 95 8.6.4 Capabilities Provided to a Payload...... 96 9 PROPULSION ...... 97 9.1 Requirements and Design Drivers...... 97 9.2 Sub-System Description...... 98 9.2.1 General...... 98

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9.2.2 Technology Trade-off...... 98 9.2.3 Generic Sub-System Design ...... 100 9.2.4 Sub-System Interfaces...... 102 9.3 Module Descriptions...... 102 9.3.1 Module A (Propellant Tank)...... 103 9.3.2 Module B (Cold Gas Generator)...... 107 9.3.3 Module C (Single Solid Propellant Thruster)...... 110 9.3.4 Module D (Single Mono-Propellant Thruster Module)...... 112 9.3.5 Module E (Single Cold Gas Thruster) & Module F (Triple Cold Gas Thruster) .....115 9.3.6 Module G (MEMS Cold Gas Thruster Pod)...... 119 9.4 Scenario Study Case 1 (LEO SSO, 3-Axis Stabilised, Without Propulsion) ...... 121 9.5 Scenario Study Case 2 (LEO SSO, 3-Axis Stabilised, With Propulsion) ...... 121 9.5.1 Design Drivers...... 121 9.5.2 Module Selection (Justification)...... 121 9.5.3 Mass/Power Budget...... 121 9.5.4 Capabilities Provided to a Payload ...... 122 9.6 Scenario Study Case 3 (GTO, Inertial Pointing Observatory) ...... 122 9.6.1 Design Drivers...... 122 9.6.2 Module Selection (Justification)...... 122 9.6.3 Mass/Power Budget...... 123 9.6.4 Capabilities Provided to a Payload ...... 123 10 POWER...... 125 10.1 Requirements and Design Drivers...... 125 10.1.1 Power Subsystem Modularization ...... 125 10.2 Sub-System Description ...... 126 10.2.1 Interfaces...... 128 10.3 Module Descriptions...... 128 10.3.1 Module 1: Solar Panel...... 128 10.3.2 Module 2: Solar Array Converter (MPPT)...... 129 10.3.3 Module 3: Battery ...... 131 10.4 Scenario Study Case 1: (LEO w/o Propulsion) ...... 131 10.4.1 Design Drivers...... 131 10.4.2 Module Selection (Justification)...... 132 10.4.3 Mass/Power Budget...... 136 10.4.4 Capabilities Provided to a Payload ...... 136 10.5 Scenario Study Case 2 (LEO With Propulsion) ...... 136 10.6 Scenario Study Case 3: GTO Mission...... 137 10.6.1 Design Drivers...... 137 10.6.2 Module Selection (Justification)...... 137 10.6.3 Mass/Power Budget...... 139 10.6.4 Capabilities Provided to a Payload ...... 139 11 AOCS...... 141 11.1 Requirements and Design Drivers...... 141 11.1.1 Attitude ...... 141

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11.1.2 Pointing Requirements ...... 141 11.1.3 Orbital Scenario Compatibility...... 141 11.1.4 Target Budgets...... 141 11.1.5 Assumptions ...... 142 11.2 Sub-System Description ...... 142 11.2.1 Interfaces ...... 142 11.3 Module Descriptions ...... 143 11.3.1 Module 1 (Digital Sun Sensor)...... 143 11.3.2 Module 2 (Star Tracker) ...... 144 11.3.3 Module 3 (Coarse Three Axis Rate Sensor (Gyro) Package)...... 145 11.3.4 Module 4 (Three axis Magnetometer) ...... 146 11.3.5 Module 5 (Navigation Camera)...... 147 11.3.6 Module 6 (GNSS)...... 148 11.3.7 Module 7 (Three Axis Reaction Wheel Package) ...... 148 11.3.8 Module 8 (Three Axis Magnetorquer Package) ...... 149 11.3.9 Module 9 (Modular AOCS Software) ...... 151 11.4 Scenario Study Case 1 (LEO – w/o Propulsion) ...... 152 11.4.1 Design Drivers...... 152 11.4.2 Module Selection (Justification)...... 152 11.4.3 Mass/Power Budget...... 153 11.4.4 Capabilities Provided to a Payload...... 153 11.5 Scenario Study Case 2 (LEO with Propulsion) ...... 155 11.5.1 Design Drivers...... 155 11.5.2 Module Selection (Justification)...... 155 11.5.3 Mass/Power Budget...... 155 11.5.4 Capabilities Provided to a Payload...... 155 11.6 Scenario Study Case 3 (GTO) ...... 155 11.6.1 Design Drivers...... 155 11.6.2 Module Selection (Justification)...... 156 11.6.3 Mass/Power Budget...... 156 11.6.4 Capabilities Provided to a Payload...... 156 11.7 Telemetry & Data Requirements...... 158 12 DATA HANDLING ...... 161 12.1 Requirements and Design Drivers...... 164 12.1.1 Budgets ...... 164 12.2 Sub-System Description ...... 165 12.2.1 Interfaces ...... 165 12.3 Module Descriptions ...... 165 12.3.1 Module 1 – Distribution & Control Module (DCM)...... 166 12.3.2 Module 2 – Remote Terminal Interface ASIC ...... 172 12.4 Scenario Study Case 1 (LEO w/o Propulsion) ...... 175 12.5 Scenario Study Case 2 (LEO with Propulsion) ...... 175 12.6 Scenario Study Case 3 (GTO) ...... 175 13 TELECOMMUNICATIONS...... 177

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13.1 Requirements and Design Drivers...... 177 13.2 Sub-System Description ...... 177 13.2.1 Modularity and Scalability...... 178 13.2.2 Interfaces...... 178 13.2.3 Modules List ...... 179 13.3 Module Descriptions...... 180 13.3.1 Module 1 (DHU Modem) ...... 180 13.3.2 Module 2 (UP/DW Converter) ...... 183 13.3.3 Module 3 (S-band LGA)...... 185 13.4 Scenario Study Case 1 (LEO w/o Propulsion) ...... 186 13.4.1 Design Drivers...... 186 13.4.2 Module Selection...... 186 13.4.3 Mass/Power Budget...... 186 13.4.4 Capabilities Provided to a Payload ...... 187 13.4.5 Scalability ...... 188 13.5 Scenario Study Case 2 (LEO with Propulsion) ...... 189 13.6 Scenario Study Case 3 (GTO) ...... 189 13.6.1 Design Drivers...... 189 13.6.2 Module Selection...... 189 13.6.3 Mass/Power Budget...... 189 13.6.4 Capabilities Provided to a Payload ...... 190 13.6.5 Scalability ...... 191 13.7 Other Solutions and Trade-Offs...... 192 13.7.1 X-Band vs. S-Band ...... 192 13.7.2 Mobile Phone Based Transponder...... 194 13.7.3 Multifunctional Distributed Antennas ...... 195 13.7.4 Electronics For Distributed Antennas...... 196 14 THERMAL ...... 197 14.1 Requirements and Design Drivers...... 197 14.2 Sub-System Description ...... 197 14.2.1 Interfaces...... 197 14.3 Module Descriptions...... 197 14.3.1 Module 1 (Black Paint)...... 197 14.3.2 Module 2 (MiSER – Miniature Satellite Energy Regulating Radiator)...... 198 14.3.3 Module 3 (Thin Plate Heat Switch) ...... 200 14.3.4 Module 4 (Heater Line – 2 Kapton Heaters + 1 Sensor) ...... 201 14.3.5 Module 5 (Heat Pipe)...... 205 14.3.6 Module 6 (MLI Blanket)...... 206 14.3.7 Module 7 (Heater With Embedded Temperature Sensor) ...... 208 14.4 Scenario Study Case 1 (LEO Scenario)...... 209 14.4.1 Design Drivers...... 210 14.4.2 Module Selection (Justification)...... 211 14.4.3 Mass/Power Budget...... 212 14.5 Scenario Study Case 2 (LEO with Propulsion) ...... 212

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14.6 Scenario Study Case 3 (GTO Scenario) ...... 212 14.6.1 Design Drivers...... 213 14.6.2 Module Selection (Justification)...... 213 14.6.3 Mass/Power Budget...... 215 15 GROUND SEGMENT & OPERATIONS...... 217 15.1 Requirements and Design Drivers...... 217 15.2 Assumptions and Trade-Offs...... 217 15.2.1 Ground Station Selection Trade ...... 217 15.2.2 Network Trade...... 217 15.2.3 Concept Trade...... 217 15.2.4 Frequency Trade ...... 218 15.2.5 Scalability Trade...... 220 15.2.6 Ground Station Equipment Trade...... 221 15.3 Baseline Design ...... 221 15.4 Options ...... 222 15.5 Technology Requirements...... 222 16 DEVELOPMENT RISK ...... 223 16.1 The Risk Assessment Process ...... 223 16.2 Risk Management Policy...... 223 16.3 Dependability and Safety Strategy ...... 224 16.4 Module Development Success Criteria ...... 224 16.5 Module Development Severity Categorization ...... 225 16.6 Risk Index...... 226 16.7 Top Risk Log...... 226 16.8 Risk Log Conclusions...... 235 17 PROGRAMMATICS ...... 237 17.1 Requirements and Design Drivers...... 237 17.2 Assumptions and Trade-Offs...... 237 17.3 Technology Development ...... 237 17.4 Model Philosophy...... 238 17.4.1 Case 1: LEO...... 238 17.4.2 Case 2: GTO ...... 240 17.5 Options ...... 240 17.6 Spacecraft Level Test Matrix ...... 240 17.6.1 Environmental Tests...... 241 17.7 Subsystem TRL ...... 242 17.8 Schedule ...... 244 17.9 Summary and Conclusions ...... 244 18 COST...... 245 18.1 Class of Estimate ...... 245 18.2 Cost Estimate Methodology ...... 245

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18.3 Scope of Estimate ...... 246 18.3.1 Individual Modules...... 246 18.3.2 LEO Mission Scenario (Option 1, 2, and 3) ...... 246 18.4 Main Assumptions...... 246 18.4.1 General Cost Assumptions...... 246 18.4.2 Technology Development Assumptions...... 247 18.4.3 Modules and LEO Mission Schedule Assumptions ...... 248 18.4.4 Ground Segment and Operations Cost Assumptions...... 248 18.4.5 ESA Internal Costs Assumptions...... 249 18.5 Technology Readiness Level Definition ...... 249 18.6 Cost Risk/Opportunity...... 249 18.6.1 Definition and Background...... 249 18.6.2 Cost Risk/Opportunity Specific Assumptions ...... 250 18.7 Conclusions and Recommendations...... 252 18.7.1 Price Growth Phenomenon and Proposed Containment Measures ...... 253 19 CONCLUSIONS ...... 255 19.1 Satisfaction of Requirements...... 255 19.1.1 Study Objectives...... 255 19.2 Final Considerations ...... 256 20 REFERENCES...... 257 21 ACRONYMS ...... 259

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1 INTRODUCTION 1.1 Background In recent years, there has been a large increase in the number of nano- and pico-satellite missions being launched. Their smaller-scale and lower-cost implementation has proved particularly well- suited to university-led research and training, as well as for enabling the test flight of some limited new technologies. Furthermore, attempts have already been made to establish configurable spacecraft ‘kits’ and the resulting cube-sat initiative has seen a very large uptake by universities and institutes. Such ‘kits’ effectively hope to enable highly modular, easily reconfigurable and yet highly capable platforms capable of supporting a wide range of payloads in a wide range of mission scenarios - however, to date, all of these s/c have made copious use of COTS electrical and mechanical components which has severely limited the lifetime. They have also not yet made full use of miniaturisation which has led to the equally severe limitations on the achievable performances of the platforms. Despite these points, some of these spacecraft have been successfully demonstrated and shown to be capable of supporting useful miniaturised payloads (see Section 3.1). 1.2 Scope This Nanosat feasibility assessment study was an ESA Internal Study Activity, requested by D- TEC, and performed under the funding of the General Studies Program. Its aim was to help the assessment of the current European state-of-the-art, the technological gaps and development needs required to realise a highly modular, easily reconfigurable and yet highly competent nanosatellite platform, capable of supporting a wide range of miniaturised payloads in a wide range of mission scenarios, for launch within ten years (2018). The study was carried out by an interdisciplinary team of specialists from ESTEC and ESOC in 8 sessions, starting with a Kick-Off on 10 December 2008 and finishing with an Internal Final Presentation on 30 January 2009. A second IFP was also held on 26 February 2009 to interested parties external to the immediate study team. The first part of the study sought to identify and define the ‘kit-of-parts’ (i.e. modules) required by each subsystem to successfully provide the platform with the requested configurability and flexibility capabilities as outlined above, as well as the miniaturisation potential of each subsystem. Subsystem presentations were used to cross-fertilize miniaturisation ideas and also to challenge and eventually re-define the traditional interfaces between sub-systems to facilitate modularity and further minaturisation. The required modules for each subsystem were then defined. The second part of the study consisted of the preliminary analysis of three case study scenarios, to demonstrate the concept and the potential services that could be provided by a payload. These case studies are concieved to provide a wide spectrum of possible application examples with different mission and environmental requirements. For each scenario, to ensure that the list of modules identified in the previous study part were sufficient and suitable, and to assess the potential services that could be provided to a payload for each scenario a preliminary, high-level, mission and platform feasibility assessment exercise has been carried out.

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In additon to the results of the above mentioned preliminary modules’ and applications’ technical assessment exercise, the outcome of the study also includes information on the development plan required to be put in place to fully realise and best exploit the nanosatellite concept in the immediate future and an assessment of the technological and programmatics risks associated to the modules’ deveolpment. A preliminary cost analysis, covering both modules’ development and nanosatellite missions’ implementations, has also been completed, to provide effort assessment information. 1.3 Document Structure The layout of this report of the study results can be seen in the Table of Contents. The Executive Summary chapter provides an overview of the study; details of each domain addressed in the study are contained in specific chapters. Due to different document distribution requirements, only cost assumptions excluding figures are given in this report. The complete costing information is published in a separate document CDF- 84(B).

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2 EXECUTIVE SUMMARY 2.1 Study Flow The Nanosat ESTEC Concurrent Design Facility (CDF) study was an ESA Internal Study Activity, requested by D-TEC, and performed under the funding of the General Studies Program to assess the usefulness and viability of a highly reconfigurable nano-satellite platform, and the design efforts required to realise this within a ten year timeframe. The study was carried out by an interdisciplinary team of specialists from ESTEC and ESOC in eight sessions, starting with a Kick-Off on 10 December 2008 and finishing with an Internal Final Presentation (IFP) on 30 January 2009. A second IFP was held on 26 February 2009 to interested parties external to the immediate study team. 2.2 Study Objectives • To produce a costed development plan to realise a nano-spacecraft for launch in 2018 and to assess the availability of technologies and units/modules within five years of development. • Preliminary feasibility assessment on how far we can reduce the size of a highly modular, multipurpose platform, capable of providing sufficient onboard resources to a miniaturised payload in different application scenarios, when applying disruptive technologies and extreme miniaturisation to all subsystems. • Payload design, launcher analysis and adapter and dispenser’s design was outside the frame of this Study. The goal is to maximise the "services" the platform can provide to a black box payload, while still maintaining the platform mass envelope. • To provide substantial inputs (technical feasibility, envisaged achievable results and limitations, required associated development effort and roadmap, cost, risk and programmatics assumptions) to the formulation of an Agency coordinated view on nanosatellite platforms and on their possible fields of application. 2.3 Requirements and Design Drivers 2.3.1 System Requirements Due to the particular approach to be followed for this CDF study, no specific mission, payload or application was applicable. No pre-defined mission requirements were hence available as study input, to derive proper system requirements for the design iterations. During the study definition phase, it was therefore decided to elaborate on preliminary assumptions and to derive information on the performances for a highly modular, multipurpose platform, capable of providing sufficient onboard resources to a miniaturised payload in different application scenarios. As the payload design was outside the frame of this Study, it was decided to propose a set of generic platform requirements that could maximise the "services" the platform can provide to a black box payload, with the initial system assumptions being: • Overall Platform Mass: 10 kg (Goal) / 20 kg (Maximum Allowed) • Generated Power: minimum 30 W

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• P/L Mass: 5 kg (Minimum) to 10 kg (Maximum) • Power Available to P/L (Nominal): 10 W (Minimum) to 20 W (Maximum) • Power Available to P/L (Standby/Eclipse): 5 W (Minimum) to 10 W (Maximum) • AOCS Requirements: o LEO Scenario: - Pointing stability: Min. 6 arcmin APE, 1 arcmin RPE (100seconds), Max. 1 arcmin APE, 10 arcsec RPE (100seconds), - Agility: Min. 0.5 deg/min, Max. 5 deg/min. o GTO Scenario: - Pointing stability: Min. 6 arcmin APE, 1 arcmin RPE (100seconds), Max. 1 arcmin APE, 10 arcsec RPE (100seconds), - Agility: Min. 0.5 deg/min, Max. 5 deg/min. • Data storage: minimum 32 GB • Data downlink: 2 Mbps (Assuming ~3.4 m dish on Ground) • Lifetime: based on 40 KRad at component level (TBD for S/A) – i.e. radiation hard components needed • ∆v capability: Minimum TBD m/s, Maximum TBD m/s • Redundancy Approach: Single String (no redundancy Æ redundancy built in at mission level) • Margin Philosophy: Margins at System Level only • The Spacecraft platform shall consist of a kit of interchangeable modules to allow rapid ‘missionisation’ and AIT o Platform adaptability shall be achieved via the selection of the appropriate modules. The modules themselves shall be considered non-adaptable (i.e. 100% recurring) o Modular onboard software to be configured for each mission • Reconfiguration time – 3 months • MAIT time – 3 months • First launch 2018. 2.3.2 Design Drivers • The platform will need to have a low recurring cost and be readily configurable for a large range of potential missions and payloads. It is expected that, to achieve this, the platform design concept needs to be mostly single string (redundancy to be provided at mission level if needed) • Modularity of design is expected to be essential to ensure the configurability of the system to different missions. 2.4 Study Approach The study was tackled in two parts: • The first part of the study sought to identify and define the ‘kit-of-parts’ (i.e. modules) required by each subsystem to successfully offer the platform configurability and

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performance capabilities as requested above, as well as to define the miniaturisation potential of each subsystem. • The second part of the study consisted of three study case scenarios, which were used to ensure that no key modules were missing, and to assess the potential services that could be provided to a payload for each scenario. 2.5 Technical Conclusions 2.5.1 Baseline Configuration The following sections provide a general overview of the NanoSat baseline configuration (non- mission specific) and highlight some generic properties which shall be applicable to all application cases: • Individual modules are 100% recurrent, plug-and-play approach • Bus architecture for power and data interfaces to all modules: Power control bus, Spacewire bus and transducer bus • Modules use a generic ASIC built on standard IPcores to support the bus interfaces or integrate these IP cores in their own ASIC • A high degree of integration between Power, DHS and Comms S/S has been proposed (see Figure 2-1) - this results in having an avionics core that manages all the “survival functions” of the S/C and could be adapted to any S/C size.

UP/DW Converter

Antennas

Tank STR DCM

MTQ DSS

Thrusters

Payload Heater/ temp sensors Figure 2-1: NanoSat DHS / Power / Communications interface • Multi (4) core > 120 MHz processor • NVRAM > 16 GiB, RAM 4 GiB (SRAM preferred) • Digital modem for Communication’s subsystem placed on DHU ASIC • Bus Voltage = 8V nominal, battery regulated • Li-Ion Battery • Thin film solar cells w/ MPPT converters • S-band used for TT&C and data downlink (baseline) – X-band option

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• Main structure composed of few base elements (see Figure 2-2) (standard panels).

Figure 2-2: NanoSat main structural elements • Main modular elements: o Wall Panels (full size panels with embedded solar cells + MPPT) o Half Wall Panels (with preform cut-outs for sensors + payload) o Top Panel (custom) o Bottom Panel (custom) o Panel Brackets/Interfaces • Minimal complexity • Not mission specific, sized for maximum load case (20 kg with only 4 panels) • Equipped as mission required (i.e. see Figure 2-3 for examples)

Figure 2-3: Multi-mission configurability • Multi-mission suitable o Predefined equipment mounting hole pattern (standard to all modules)

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o Removable cut-outs o Integrated support for panel attachment,hinges, hold down mechanisms etc. • Power (SA) sizing element o Min. 340mm x 230mm panel. 2.5.2 List of Modules The following comprises a comprehensive list of all the subsystem modules recommended by each discipline expert. This forms the “kit–of-parts” from which each of the investigated missions is assembled. Details of which modules belong to which option can be found in Table 5-1, Table 5-4 and Table 5-6:

Mass (kg) per unit Structure Side Panel 0.16 Panel Brackets 0.03 Top Panel 0.10 Harness 0.60 Half Panels 0.08 Bottom Panel 0.15 Deployable Panel 0.16 Horizontal Brackets - Long 0.02 Horizontal Brackets - Short 0.01 Thermal Control Black paint 0.30 MiSER (Miniature Satellite Energy Regulating Radiator) 0.15 Thin Plate Heat Switch 0.02 Digital Heater line (2 heaters+1 sensor) 0.09 Heat pipe 0.20 MLI blankets (mass is per m2) 0.50 Digital Heater with embedded sensor 0.05 Mechanisms S/A Deployment Mechanism SDM 0.09 Deorbit Deployment Mechanism (DDM) 0.29 Hold-Down and Deployment Mechanism (HDRM) 0.04 Nano-Terminator Deorbit Module (NTDM) 0.06 Communications Lightweight S-band LGA 0.02 S-band 1W UP/DW converter 0.90 S-band 2W UP/DW converter 0.90 S-band 4W UP/DW converter 0.90 S-band 8W UP/DW converter 0.90 Lightweight X-band LGA 0.10 X-band- EO 1W UP/DW converter 0.90 X-band- EO 2W UP/DW converter 0.90

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X-band- EO 4W UP/DW converter 0.90 X-band- EO 8W UP/DW converter 0.90 X-band- SR 1W UP/DW converter 0.90 X-band- SR 2W UP/DW converter 0.90 X-band- SR 4W UP/DW converter 0.90 X-band- SR 8W UP/DW converter 0.90 Cables 0.05 (Option) Mobile phone based transponder 0.30 Multifunctional distributed antennas 0.01 Electronics for distributed antennas 0.15 Data Handling DCU 0.70 Spacewire interface 0.05 TB interface 0.10 RF electronics 0.10 AOCS Digital Sun Sensor 0.06 Star Tracker 0.18 Coarse Three Axis Gyro 0.11 3 axis Magnetometer 0.05 Navigation Camera 0.18 GNSS 0.18 Three Reaction Wheel Package 0.46 Three Axis Magnetorquer 0.10 Propulsion Propellant Tank Module (A) Tank 1.25 Propellant Tank Module (A) Fill/Vent Valve 0.07 Propellant Tank Module (A) Fill/Drain Valve 0.07 Propellant Tank Module (A) Filter 0.10 Propellant Tank Module (A) Helium/Pressurant 0.00 Cold Gas Generator Module (B) Tank Assembly 0.13 Cold Gas Generator Module (B) Pressure Transducer 0.08 Cold Gas Generator Module (B) Electronics 0.20 Solid Rocket Module (C) Motor 0.48 Solid Rocket Module (C) Nozzle 0.10 Solid Rocket Module (C) Ignition/Electronics 0.10 1 Thruster Monoprop Module (D) Thruster 0.20 1 Thruster Monoprop Module (D) Latch Valve 0.65 1 Thruster Monoprop Module (D) Electronics 0.10 1 Thruster Monoprop Module (D) Piping 0.05 1 Thruster Monoprop Module (D) Wiring 0.05 1 Thruster Butane Module (E) Thruster 0.08 1 Thruster Butane Module (E) Vaporiser 0.05 1 Thruster Butane Module (E) Electronics 0.10 1 Thruster Butane Module (E) Piping 0.05 1 Thruster Butane Module (E) Wiring 0.05 3 Thruster Butane Module (F) Thruster 0.23

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3 Thruster Butane Module (F) Vaporiser 0.05 3 Thruster Butane Module (F) Electronics 0.15 3 Thruster Butane Module (F) Piping 0.15 3 Thruster Butane Module (F) Wiring 0.15 4 Thruster MEMS Nitrogen CG Module (G) Thuster Pod 0.12 4 Thruster MEMS Nitrogen CG Module (G) P Transducer 0.00 4 Thruster MEMS Nitrogen CG Module (G) Electronics 0.20 4 Thruster MEMS Nitrogen CG Module (G) Piping 0.05 4 Thruster MEMS Nitrogen CG Module (G) Wiring 0.00 Power Battery 0.27 Solar Array 0.18 PCDU 0.12 Table 2-1: Complete NanoSat module list A programmatic assessment was then carried out on this list of modules to analyse the feasibility of attaining the required TRL for each module in time to allow the launch of the first NanoSat in 2018. This programmatic development plan for each module is as given in Table 2-2.

2008 2009 2010 2011 2012 2013 2014 2015 2016 2017 2018 AOCS Digital Sun Sensor 3689 Star Tracker 23679 Gyro 56 8 9 Earth Sensor 137 GNSS Receiver 46 8 9 Magnetometer 468 9 Magnetorquer 46 8 9 Reaction Wheel 34689 Navigation camera 23679

Antennas/Comms Lightweight S-band antenna 689 Lightweight X-band antenna 2 3 689 Multifunctional distributed antenna system 124689 Electronics for distributed antenna system 12 4 6 89 UP/DW converter efficient power amp. 3689 Mobile phone based transponder 234689

DHS Control Distribution Unit 2489 General purpose Interface ASIC 358 DCM (SoC ASIC) -System on a chip-

Power Solar Array 23456 Battery Pack 23456789 Power Conditioning 45678

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2008 2009 2010 2011 2012 2013 2014 2015 2016 2017 2018 Propulsion Propellant Tank Module (A) 56778 Cold Gas Generator Module (B) 456678 Solid Propellant Thruster Module (C) 2334556678 Single Thruster Monopropellant Module (D) 34556678 Single Thruster Butane Module (E) 34556678 Three Thruster Butane Module (F) 34556678 Four Thruster MEMS Nitrogen Module (G) 34556678

Structure Conventional Structure 79 Innovative Structure 2 3456789 Harness based on Nanotubes 12456 Conventional Harness 79

Mechanisms S/A Deployment Mechanism (SDM) 456 8 Deorbit Deployment Mechanism (DDM) 456 8 Hold down and Release Mechanism (HDRM) 28 Nano-Terminator Deorbit Module (NTDM) 28

Thermal Black paint 8 MiSER (Miniature Satellite Energy Regulating Radiator) 8 Thin Plate Heat Switch 8 Heater line (2 heaters+1 sensor) 8 Heat pipe 8 MLI blankets 8 Table 2-2: Programmatic development plan for subsystem modules This analysis highlighted the following issues: • The programmatic target (2018) used in the Study for the first nanosat mission launch seems in line with most of the modules’ development time schedules. It must be however highlighted that the development time assumptions used in this Study are more optimistic than those described in the “ European Space Technology Master Plan”. • Heavy investments in technology development, design, manufacturing and testing shall be put in place for the identified with lower TRL in 2009 and for Interface IP core plus ASIC development, to ensure their timely readiness for the realisation of a nanosatellite platform in the given timeframe. 2.6 Test Mission Scenarios Three study cases application scenarios were used to demonstrate and evaluate the potential services that the NanoSat modules, when combined, could provide to a payload: • LEO SSO (circular, 600 km, 06:00/09:00/12:00 LTAN) 3-Axes Stab. w/o Propulsion – Assuming Earth Pointing Dummy P/L Package • LEO SSO (circular, 600 km, 06:00/09:00/12:00 LTAN) 3-Axes Stab. with Propulsion - Assuming Earth Pointing Dummy P/L Package • GTO (600Rp/35943Ra, 6deg Incl) - Assuming Inertial Pointing (Observatory) Dummy P/L Package 2.6.1 Scenario 1 – LEO SSO (w/o Propulsion) Scenario assumptions: • LEO 600 km altitude circular SSO

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o (Link Budget to consider 800 km altitude, to get a worst case analysis for this S/S) • 12:00 LTAN (to size for maximum eclipse - ~35 mins) • 3-Axes Stabilised - Nadir Pointing Spacecraft (with pointing requirements given above) • No onboard propulsion • P/L Input Power: 5 W (continuous) or 10 W (50% Duty Cycle) • 2 Mbps Data rate • Earth Pointing Dummy P/L Package. Based on these assumptions, modules were selected from Table 2-1 and combined to give a platform with potential performance characteristics as follows:

Platform S/C mass Dry mass 9.12 kg (incl. 20% system margin) Wet mass 9.12 kg Available payload 22 L volume (230 x 280 x 340 mm) Available payload 10.88 kg mass (assuming total S/C launch mass of 20 kg) Available payload 10 W (nominal & eclipse, with 50% duty cycle) power Communications 4.5 Gb download per day * AOCS performance Determination APE [arcmin] 0.4 RPE [arcsec/100] 4 Pointing APE [arcmin] 1 RPE [arcsec/100] 10 Slewing 90 deg in < 100 sec * For a link budget at 800 km altitude with 30 min of contact per day at Redu Table 2-3: NanoSat performance for Test Scenario 1 (LEO w/o propulsion) 2.6.2 Scenario 2 – LEO SSO (w/ Propulsion) The second study case application was similar to Scenario 1, albeit with onboard propulsion included. The scenario definition is: • LEO 600 km altitude circular SSO (10:30 LTAN – typical range for many EO missions) o (Link Budget to consider 800 km altitude) • Onboard propulsion for 1 DoF control o (max. 5 kg total propulsion S/S mass allocated) • 3-Axes Stabilised - Nadir Pointing Spacecraft

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• P/L Input Power: 5 W (continuous) or 10 W (50% Duty Cycle) • 2 Mbps Data rate. Based on these assumptions, modules were selected from Table 2-1 and combined to give a platform with potential performance characteristics as follows:

Platform S/C mass Dry mass 11.30 kg (incl. 20% system margin) Wet mass 13.29 kg Available payload 8 L volume (250 x 280 x 120 mm) Available payload 6.71 kg mass (assuming total S/C launch mass of 20 kg) Available payload 10 W (nominal & eclipse, with 50% duty cycle) power Communications 4.5 Gb download per day *

AOCS performance Determination APE [arcmin] 0.4 RPE [arcsec/100] 4 Pointing APE [arcmin] 1 RPE [arcsec/100] 10 Slewing 90 deg in < 100 sec Available ∆V 1 DoF 62.50 m/s (using Cold Gas Butane, Isp = 62 s) * For a link budget at 800 km altitude with 30 min of contact per day at Redu Table 2-4: NanoSat performance for Test Scenario 2 (LEO w/ propulsion) 2.6.3 Scenario 3 – GTO Scenario assumptions: • Injection orbit: 250 x 35143 km o with 6º inclination wrt Equator • Final orbit (example): 600 x 35143 km o 10.4 hrs period (max eclipse duration 2.2 hrs) • Onboard propulsion for 3 DoF control and momentum offloading o (max. 5 kg wet mass allocated) • 3-Axes Stabilised – Inertial/Sun Pointing Spacecraft • P/L Input Power: 10 W continuous • 2 Mbps Datarate.

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Based on these assumptions, modules were selected from Table 2-1 and combined to give a platform with potential performance characteristics as follows:

Platform S/C mass Dry mass 16.48 kg (incl. 20% margin) Wet mass 18.47 kg Available payload 8 L volume (250 x 280 x 120 mm) Available payload 1.5 kg (approx) mass Available payload 10 W (continuous) power Communications 1.1 Gb download per day * AOCS performance Determination APE [arcmin] 0.4 RPE [arcsec/100] 4 Pointing APE [arcmin] 1 RPE [arcsec/100] 10 Slewing 90 deg in < 100 sec Available ∆V 3 DoF 62.50 m/s (using Cold Gas Butane, Isp = 62 s) 1 DoF (de- 50 m/s orbiting) (using solid propellant, Isp = 260 s) * For a link budget considering use of VCM at Malindi and a transmission window between 3000 and 11000 km (37 min/day) Table 2-5: NanoSat performance for Test Scenario 3 (GTO)

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3 STUDY OBJECTIVES AND APPROACH 3.1 Miniature Satellite Background A large number of nano-sats and pico-sats have already been designed and launched. Attempts have also already been made to establish configurable s/c ‘kits’ and the resulting cube-sat initiative has seen a very large uptake by universities and institutes. These cube-sats provide a cost effective means of obtaining hands-on training and learning as well as enabling the test flight of some limited new technologies. However, to date, all of these s/c have made copious use of COTS electrical and mechanical components which has severely limited the lifetime. They have also not yet made full use of miniaturisation which has led to the equally severe limitations on the achievable performances of the platforms. Despite these points, some of these spacecraft have managed impressive performances and been shown to be capable of supporting useful payloads. Whereas nano- and pico-sats have been designed and launched from all over the world, there appears to be a clear lead in the area coming from USA and Asia. It is also clear that NASA and the USAF provide the funding for the majority of this class of s/c in the USA. This compares markedly with Europe where most activities in this area are by individual universities. It is also clear that the 5 to 25 Kg class of s/c have already seen a number of missions with significantly useful payloads (e.g. 3CS, Hausat-2, Tubsat, nano-jasmine…). This has been possible even without significant efforts in the miniaturisation and radiation hardening of the spacecraft components. Table 3-1 shows some examples of recent miniature satellite missions and their payloads.

Mission Country Organisation Mass (Class) Payload 3CS USA University of 15 kg (Nano) Stereoscopic Colorado, US imaging, Air Force Formation flying (launcher failure) HAUSAT -2 Korea Hankuk 25 kg (Sub-Micro) Animal Tracking Aviation Plasma probe University and MOST NanoSpace-1 China Qinghau 25 kg (Sub-Micro) Technology Satellite demonstrations Technologies SNAP-1 UK SSTL 6.5 kg (Nano) ST5 USA NASA 25 kg (Sub-Micro) Magnetic field measurements Table 3-1: Examples of recent miniature spacecraft missions

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3.2 Modularity – The Cubesat Example The concept of a modular s/c bus, rapidly configurable from a ‘kit of parts’ is not new. The Cubesat initiative, started by Professor Twiggs of Stanford University, has successfully used the same approach to make pico-satellites affordable and available to universities as teaching and research tools. The success of this initiative with respect to its goals, makes it interesting to summarise here as background information and a demonstration that the availability of a kit of parts is an effective means of reducing the recurring cost while – if the modularity level is properly selected – not unduly limiting the configuration or capability possibilities. The Cubesat concept provides a catalogue of module options that are all interoperable to allow the construction of a 10*10*10 cm, 1 kg satellite. This spacecraft has been broken down into modules covering software, structure, main electronics board, memory options, processor options, power supply, attitude control, communication and development support tooling. By selection from this kit of parts it is possible to build a wide variety of configurations – even up to 30*10*10 cm – in a fast, efficient and low cost way. To illustrate the concept further, Table 3-2 below gives examples of some of the kit modules available for several of the sub-systems.

Subsystem Modules Structure Cover plate Base plate ½ size chassis wall Full size chassis wall Double size chassis wall Hinge kit Electronics mounting kit Solar Panel mounting kit

Power Power conditioning board (supports up to X batteries and Y Solar array panels) Cubesat Battery module Solar Array panel

Comms MHX-920A Transceiver Module MHX-2400 Transceiver Module MHX-2420 Transceiver Module MHX-2420-SL Transceiver Module Table 3-2: Examples of Cubesat kit modules and options

Each of these modules are provided with well defined and ‘standard’ interfaces, user manuals and full CAD models to further ease the custom configuration.

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Figure 3-1: A Cubesat structural assembly example using off the shelf structural element modules 3.3 Study Objectives & Trade-Offs • To produce a costed development plan to realise a nano-spacecraft for launch in 2018 and to assess the availability of technologies and units/modules within 5 years of development • Preliminary feasibility assessment on how far the size of a highly modular, multipurpose platform, capable of providing sufficient onboard resources can be reduced to a miniaturised payload in different application scenarios, when applying disruptive technologies and extreme miniaturisation to all subsystems • The payload design was outside the frame of this Study. The goal is to maximise the "services" the platform can provide to this black box payload, while still maintaining a fixed platform mass envelope • The platform will need to have a low recurring cost and be readily configurable for a large range of potential missions and payloads. It is expected that, to achieve this, the platform design concept needs to be single string (redundancy to be provided at mission level if needed) • Modularity of design is expected to be essential to ensure the configurability of the system to different missions • The investigation to be based on a development roadmap assuming launch of the first nano-platform of this kind in ten years from now • Launcher analysis, adapter and dispenser’s design was outside the frame of this study • Provision of substantial inputs (technical feasibility, envisaged achievable results and limitations, required associated development effort and roadmap, cost, risk and programmatics assumptions) to the formulation of an Agency coordinated view on nanosatellite platforms and on their possible fields of application. 3.4 Study Starting Assumptions and Generic Platform Requirements • Spacecraft must be rad hard to 40 kRad • Single string only – redundancy to be achieved at mission level • Mission re-configuration time of 3 months

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• Mission AIT time of 3 months • Mission lifetime targeted at 3 years minimum • First launch in 2018.

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4 MISSION ANALYSIS The mission analysis task, in the framework of the Nanosat CDF study, has been to identify and propose mission scenarios compatible with the design and applications of a Nanosat. The task comprises the following activity: • Analysis of a mission scenario in LEO not requiring propulsion • Analysis of a mission scenario in LEO requiring propulsion • Analysis of a mission scenario in GTO. Also included is the generation of products associated with the above analysis. These products are made available to the other sub-systems involved in the design process. 4.1 Requirements and Design Drivers 4.1.1 Mission Objective and Reference Mission Scenario The mission analysis shall identify possible scenarios in order to assess the feasibility of a Nanosat in the mass range of 10-20 kg. Particularly realistic mission profiles which provide interesting design cases for the s/c and the different subsystems shall be identified. More specifically the following scenarios shall be analyzed in details: • Low Earth Orbit: Sun-Synchronous • Geostationary Transfer Orbit. 4.1.2 Launch Year and Mission Duration Earliest mission launch is in 2018. The mission duration shall be 3 years or shall not exceed a maximum radiation dose equal to 40 kRad (at component level). 4.1.3 Orbit Maintenance The mission profiles identified shall provide mission options for a s/c design which does not need onboard propulsion, and options for which propulsion is necessary for orbit acquisition, maintenance or end of life de-orbiting. Particularly the LEO case shall present a propulsion free option as well as an option which requires full orbit control. No specific requirement is set for the orbit maintenance. 4.1.4 End of Life Disposal The mission and spacecraft design shall comply with the guidelines for the Space Debris mitigation which require that the s/c re-enters within 25 years after the end of the operations. 4.2 Assumptions and Trade-Offs 4.2.1 S/C Size and Mass The following assumptions have been made for the size and weight of the spacecraft: • Bus mass : 10 kg • Payload mass: 5 kg

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• Propulsion mass: 5 kg maximum • Size: from 0.3x0.3 m to 0.46x0.46 m. Based on these assumptions, the ballistic coefficient can vary between 30 kg/sqm and 70 kg/sqm. 4.2.2 LEO Option: Altitude and Maintenance Cost Sun-Synchronous orbits are flown at different altitude, typically ranging from 450 to 820 km. At low altitude the main driver for the ∆v budget is the orbit maintenance to compensate the effect of drag, which strongly perturbs the orbit to decay. At higher altitude the cost for orbit maintenance is less demanding, although the 3rd body perturbation on the inclination affects the local time, and most of the ∆v budget is taken up by the de-orbit burn – however, due to the quite low ballistic coefficient (in the order of 30 to 50 kg/sqm), high altitude SSO can still be considered in this case. No relevant differences exist in terms of coverage or eclipse duration. As the requirement in terms of the orbit control needed by the payload is not specified, typical values used for sun-synchronous orbits have been taken as reference. The following assumptions have been taken for the requirements on the orbit control: • LTAN: +/- 10 minutes control box • Groundtrack: +/- 30 km. • Mission lifetime: 3 years. As an example a comparison of orbit maintenance manoeuvres is here presented between a low and high altitude Sun-Synchronous orbit. Nominal altitude: 820 km. At this height the drag is almost negligible, but it still affects the semimajor axis of the orbit causing a drift in groundtrack. However most of the orbit maintenance is driven by the local time control, which is perturbed by the 3rd body in by changing the inclination: • Correction of injection error: 10 m/s • Target orbital position acquisition: 2.5 m/s • Ground track control: 8 m/s. There are 16 maneuvers on average every 65 days. • Local Time Control: 12 m/s (drift < 10 minutes) • De-orbiting Maneuver: 70-80 m/s. Nominal altitude: 550 km. If the nominal altitude is around 550 km, then drag has a large impact and affects both local time and groundtrack. This option is clearly very much affected by the actual drag and solar activity. • Correction of injection error: 10 m/s • Target orbital position acquisition: 2.5 m/s • Groundtrack control: 26 m/s. There is a maneuver every 42 days, with each burn being of 0.7 m/s • Local Time is consistently and automatically corrected due to the correction cycle in groundtrack. Max drift is approximately 5 minutes • No ∆v for re-entry.

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If the cost for de-orbiting for a high altitude SSO is not included in the ∆v budget, assuming that de-orbiting could be achieved with different ad hoc devices, then it can be concluded that, under the given assumptions, the Nanosat could fly any SSO altitude ranging from approximately 500 to 800 km, with a very similar ∆v budget allocation. 4.2.3 GTO The standard GTO orbit is intrinsically stable for decades. Therefore in theory the perigee could be left at its nominal altitude at 250 km. However the numerous passes in the atmosphere would affect the stability of the pointing of the payload (here assumed to be an observatory payload), and it has been decided to raise the perigee to a less disturbed altitude. This implies a trade-off between the cost of the manoeuvre, which is double as it is also necessary at the end of the mission to de-orbit, and the minimum stable altitude. Based on the ∆v cost presented in Figure 4-1, it has been assumed that the perigee will be raised up to 600 km. For the perigee raising sequence a trade-off analysis has been performed based on the different propulsion systems available. Two cases have been analyzed: a cold gas system and a monopropellant system. The monopropellant system allows doing the perigee raising manoeuvre in one go, while for the cold gas case three burns are required. The impulsive ∆v to raise the perigee from 250 km to 600 km is 35.7 m/s. The monopropellant system has the following propulsive characteristics: maximum thrust =0.5 N, Isp=216 s. The burn takes place around the apogee starting at true anomaly 178.46 deg and ending at true anomaly 181.54. Total thrust time is 1420 s. The gravity losses are 0.12 %.

The cold gas engine has a maximum thrust of 50mN, Isp=62 s. The analysis has been performed with one, two or three burns: • 1 burn: With one single burn the gravity loss is in the order of 13.2 % with a total thrust time of 4.5 hours. The real ∆v becomes: 40.6 m/s • 2 burns: Gravity losses can be kept in the order of 3%. A first burn raises the perigee to 450 km, with a burn duration of 8430 seconds, while the second burn reaches the final perigee altitude of 600 km in additionally 5900 seconds. Total actual ∆v is 36.8 m/s • 3 burns: Gravity losses are kept below 1.5 %. A first burn raises the perigee to 350 km with a total duration of 4150 s. The second brings the perigee up to 450 km in approximately 4000 seconds and the last to 600 in 6000 seconds. the total ∆v is 36.12 m/s. In all cases all burns occur at apogee under illumination conditions. These burns could be considered to happen right after the separation of the s/c and according to typical flight dynamics rule a 3 days gap between each burn (for Orbit determination) should be considered. The final orbit, in the three burns case, could be achieved in 6 days.

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Figure 4-1: Cost to change the perigee (left) and the apogee (right)

The cost of various manoeuvre options from a standard GTO (perigee at 250 km) are the following: • Transfer to GEO :1505 m/s • Reducing the perigee down to 120 km for a controlled re-entry: 14 m/s • Going to the moon: 680 m/s • More generally changing the perigee costs approximately 0.1 m/s per km • Changing the apogee costs 0.016 m/s per km • Changing the inclination would cost 28 m/s per degree. 4.2.4 De-orbiting It is assumed that for both scenarios the s/c will comply with the space debris guidelines. The s/c design shall decay within 25 years either by means of a re-entry manoeuvre at the end of the mission or by natural decay. For the SSO case, due to ballistic coefficient in the order of 30-50 kg/sqm, deorbiting occurs naturally up to 650 km within 25 years. If the ballistic coefficient is reduced down to approximately 20 kg/sqm then re-entry within 25 years can be ensured up to 700 km. A reduction of the ballistic coefficient down to 20 kg/sqm is possible by means of mechanical devices which increase the reference surface area of the s/c at the end of life. Above 700 km altitude it is probably preferable to resort to specific de-orbiting device. It is important to underline that due to the flexibility of the modular design, the Nanosat s/c is capable of flying at all the typical altitudes of the Sun-Synchronous orbits, and yet respect the guidelines on the Space Debris. For the GTO case in order to ensure a re-entry under any conditions it is necessary to lower the pericenter down to a safe value of approximately 120 km. In fact due to the precise orientation of the GTO orbit with respect to the sun at the end of the mission the 3rd body perturbation can act in different ways. Under some conditions in fact the eccentricity decrease and the perigee increases. If the perigee is not sufficiently low (i.e. around 150 km) then the re-entry might not occur. Conversely if 120 km is achieved, then re-entry occurs under any condition, in less than 10 years. The cost of lowering the perigee from 600 down to 140 km is 46.5 m/s. If 50 m/s are allocated then approximately 120 km perigee is achieved and re-entry is safe.

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4.3 Baseline Design 4.3.1 Scenario 1: SSO Orbit Without Propulsion A LEO scenario without propulsion has been envisaged, and it has been identified in a SSO mission profile. The nominal altitude is set at 600 km and different initial local times of the ascending node have been considered to assess their influence on the other subsystems’ design. As the s/c is not equipped with propulsion, the orbit is uncontrolled and affected mainly by the perturbation due to the drag, which lowers the altitude of the s/c, thus affecting the sun- synchrounous conditions which depends on the semimajor axis, the inclination and the eccentricity. This has two consequences: The first effect is on the local time of the ascending node which experiences a very large drift over the mission lifetime. Depending on the solar activity and the launch date the local time can have a drift of more than 1.5 hours over the mission lifetime. The second consequence is on the groundtrack. For such a case the concept of groundtrack repetition is not applicable as due to the lack of an active orbit control, the groundtrack will quickly drift in longitude. The lifetime, at the selected altitude, is ensured regardless of the perturbation of the drag, and re- entry occurs within few years after the end of the mission. 4.3.2 Scenario 2: SSO Orbit With Propulsion For the LEO scenario with propulsion a standard SSO at 600 km is assumed. The nominal ∆v budget is shown below. This is based on the assumption made for the orbit control requirements: • Orbit injection correction: 3 m/s to correct error in the semimajor axis of 5 km + 6.5 m/s to correct the inclination in order to meet the correct sun-synchronous conditions. In total it is 9.5 m/s of injection error correction. • Target position acquisition: 5 m/s. This ∆v allows changing the nominal angular position on the orbit. This could be used to reach an angular position different from that of the launcher or the main s/c. Alternatively this corresponds to 10 km change in semimajor axis, which could just be used to separate in altitude and not in true anomaly the Nanosat. • Ground track control: 40 m/s. This is a very generous allocation for the control of the ground track within +/- 30 km. This control at 600 km in the worst case scenario would require 9 manoeuvres per year each of approximately 1.5 m/s (the actual value is 1.1 but a generous margin has been added). A 3 years lifetime would require 40 m/s. • Inclination control 6.5 m/s: Assuming a variation of 0.05 degrees of inclination in 3 years, additionally 6.5 m/s could be budgeted for precise control of the inclination which would allow a better control on the Groundtrack at the pole and a fine tuning of the local time. • The above values are assumed as the nominal ∆v for the mission; they sum up to 61 m/s. If a larger allocation for ∆v should be available, some options could be considered which could be compatible with the current mission scenario and enhance the value of the mission or extend its duration. If 110 m/s are available the additional 50 m/s could be used to extend the mission lifetime. For example it could be envisaged after three years or even earlier to lower the altitude to obtain better resolution of the observation instrument.

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The following proposal is made: • 41.5 m/s would allow reducing the perigee to 400 km and flying on an elliptic orbit with perigee at 400 km and apogee at 600 km. Anyway this option would modify the SSO condition, which would be lost. • Alternatively it could be possible to lower the orbit down to a circular 550 km orbit, which would require 28 m/s. The ground track control would still be possible and would cost 3.4 m/s per manoeuvre every 24 days. As after the circularization burn there are still 21 m/s left of the total ∆v budget, this would be enough to extend the mission for additional 5 months. 4.3.3 Scenario 3: GTO In the scenario 3 the Nanosat is injected into a standard GTO, as a piggyback of a commercial launch. Therefore the standard parameters of a GTO midnight launch orbit have been assumed: • Perigee altitude: 250 km • Apogee altitude: 35943 km • Inclination: 6 deg • Argument of perigee: 178 deg. The value of the argument of perigee implies that regardless of the season the perigee lies approximately in the Earth equatorial plane, while the midnight launch implies that the apogee is illuminated at the beginning of the mission. Regardless of the exact launch date, during the mission lifetime the orbit is perturbed and at some time during the mission it will experience large eclipse periods. This happens when the apogee will be on the night side. Eclipses as can be seen from Figure 4-2 will be as long as 2.2 hours at some point during a 3 years mission.

Figure 4-2: Eclipse evolution for a standard GTO launched on the Spring Equinox

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Figure 4-3: Evolution of the altitude Immediately after launch the s/c is manoeuvred to raise the perigee up to 600 km. The baseline is to do this with the a Cold Gas propulsion system which may need more than one manoeuvre. This costs 36 m/s. During the mission no orbit control is required. Re-entry is ensured at the end of the mission by lowering the perigee down to 120 km with a ∆v cost of 50 m/s. This is exactly the ∆v capability that the (separate) engine module assumed for the de-orbiting manoeuvre is capable of delivering. The most interesting aspect from a s/c design point of view of this GTO scenarios are the long eclipses and the long time spent at large distance from the Earth, as can be seen in Figure 4-3. 4.4 Budgets Table 4-1 lists the ∆v budget for the required mission manoeuvres. These values can be used to compute the actual mass into the target orbit as well as the propellant, dry and wet masses.

Scenario 1 Scenario 2 Scenario 3 Launcher injection N/A 9.5 m/s 36 m/s Correction Orbit Acquisition N/A 5 m/s 0 Orbit Control N/A 47 m/s 0 De-orbit Burn N/A 0 m/s 50 m/s Total ∆V 0 m/s 61 m/s 86 m/s Table 4-1: ∆v budget

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5 SYSTEM 5.1 Purpose and Scope The aims of this chapter are twofold: • To gather and present the development activities required at system level, both to coordinate and guide the sub-system module developments and to develop the system level tools and equipment required for the design and AIT of a nano-spacecraft. Without these co-ordinating activities, the module developments run a high risk of diverging and of never being capable of being assembled into a working platform. A number of the activities are also targeted at reducing the development and AIT (and hence the cost) of any nano-satellite mission assembled from the developed modules. • To present the resultant system designs for the three considered mission test scenarios, which were used to ensure that no key modules were missed from the subsystem proposals, and to assess the potential services that could be provided to a payload for each scenario. 5.2 System Requirements Due to the particular approach to be followed for this CDF study, no specific mission or application was applicable. No pre-defined mission requirement was therefore available as study input, to derive proper system requirements for the design iterations. During the study definition phase, it was therefore decided to elaborate on the preliminary assumptions, to derive information on the performances a highly modular, multipurpose platform, capable of providing sufficient onboard resources to a miniaturised payload in different application scenarios shall have. As the payload design was outside the frame of this Study, it was decided to propose a set of generic platform requirements that could maximise the "services" the platform can provide to this black box payload, while still remaining within the envisaged mass envelope: • Overall Platform Mass: 10 kg (Goal) / 20 kg (Maximum Allowed) • Generated Power: minimum 30 W • P/L Mass: 5 kg (Minimum) to 10 kg (Maximum) • Power Available to P/L (Nominal): 10 W (Minimum) to 20 W (Maximum) • Power Available to P/L (Standby/Eclipse): 5 W (Minimum) to 10 W (Maximum) • AOCS Requirements: o LEO Scenario: - Pointing stability: Min. 6 arcmin APE, 1 arcmin RPE (100 seconds), Max. 1 arcmin APE, 10 arcsec RPE (100 seconds), - Agility: Min. 0.5 deg/min, Max. 5 deg/min. o GTO Scenario: - Pointing stability: Min. 6 arcmin APE, 1 arcmin RPE (100 seconds), Max. 1 arcmin APE, 10 arcsec RPE (100 seconds), - Agility: Min. 0.5 deg/min, Max. 5 deg/min.

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• Data storage: minimum 32GB • Data downlink: 2 Mbps (Assuming ~3.4 m dish on Ground) • Lifetime: based on 40 KRad at component level (TBD for S/A) – i.e. radiation hard components needed • ∆v capability: Minimum TBD m/s, Maximum TBD m/s • Redundancy Approach: Single String (no redundancy Æ redundancy built in at mission level) • Margin Philosophy: Margins at System Level only • The Spacecraft platform shall consist of a kit of interchangeable modules to allow rapid ‘missionisation’ and AIT o Platform adaptability shall be achieved via the selection of the appropriate modules. The modules themselves shall be considered non-adaptable (i.e. 100% recurring) o Modular onboard software to be configured for each mission • Reconfiguration time – 3 months • MAIT time – 3 months • First launch 2018. 5.3 System Design Drivers • The platform will need to have a low recurring cost and be readily configurable for a large range of potential missions and payloads. It is expected that, to achieve this, the platform design concept needs to be single string (redundancy to be provided at mission level if needed). • Modularity of design is expected to be essential to ensure the configurability of the system to different missions. 5.4 System Development Activities 5.4.1 Requirements and Design Drivers The Systems Engineering activities have the following drivers: a) Ensure a detailed set of consistent interface specifications and standards are applied to all developments (data, power, mounting, environmental conditions). b) Perform the systems engineering tasks required to provide the ‘glue’ between all modules and maintain the modularity and flexibility of the platform. c) To develop the configuration and analyses tools required to rapidly identify the ‘best fit’ set of modules needed to assemble a s/c to meet mission specific requirements and from there determine the properties and performances of the resulting spacecraft. d) To develop and verify the AIT tools and GSE required to rapidly and efficiently assemble and test a nano-satellite platform. e) To develop and validate a suitable, generic launcher interface and ‘mothercraft’ interface adaptors.

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f) To co-ordinate with the ground systems development and ensure inter-operability between the ground system and the on-board software. g) To develop and verify a modular and flexible on board data handling application software. 5.4.2 Activity/ Module List The following activities are foreseen under the Systems Engineering development heading: • Systems requirement definition and flow down • Interface standards definitions including Launcher interface development • Micro connector development and qualification • Modularity systems engineering • On-Board DHS Application Software Development • Rapid system configuration, performance and budgets tool development • System level bench test trials and AIT process development • Recurring cost assessment and optimisation. 5.5 Mission Baselines - Scenario 1 (LEO w/o Propulsion) The first test case represented a Low Earth Orbit (LEO) scenario, with a black-box, undefined payload, and no onboard propulsion. De-orbiting within 25 years is considered an essential mission requirement. The scenario definition is: • LEO 600 km altitude circular SSO o (Link Budget to consider 800 km altitude) • 12:00 LTAN (to size for maximum eclipse - ~35 mins) • 3-Axes Stabilised - Nadir Pointing Spacecraft (with pointing requirements given above) • No onboard propulsion • P/L Input Power: 5 W (continuous) or 10 W (50% Duty Cycle) • 2 Mbps Data rate • Earth Pointing Dummy P/L Package. 5.5.1 System Baseline Design 5.5.1.1 Description Taking into account the mission objectives, requirements and assumptions outlined above, the following system baseline was defined. 5.5.1.2 Module Selection / Equipment List The following list presents the module selection for the NanoSat spacecraft in Scenario 1. Note the module selection justifications are given in the individual subsystem sections, and are not repeated here.

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Element 1 - NanoSat FUNCTIONAL SUBSYSTEM nr Mass (kg) per unit Total Mass (kg) Structure 2.27 Side Panel 4 0.16 0.62 Panel Brackets 8 0.03 0.24 Top Panel 1 0.10 0.10 Harness 1 0.60 0.60 Half Panels 4 0.08 0.31 Bottom Panel 1 0.15 0.15 Horizontal Brackets - Long 8 0.02 0.16 Horizontal Brackets - Short 8 0.01 0.08 Thermal Control 0.39 Black paint 1 0.30 0.30 Heater line (2 heaters+1 sensor) 1 0.09 0.09 Mechanisms 0.06 Nano-Terminator Deorbit Module (NTDM) ( 1 0.06 0.06 Communications 0.99 Light S-band LGA 2 0.02 0.04 Cables 1 0.05 0.05 UP/DW converter 2W 1 0.90 0.90 Data Handling 0.95 DCU 1 0.70 0.70 Spacewire interface 1 0.05 0.05 TB interface 1 0.10 0.10 RF electronics 1 0.10 0.10 AOCS 1.19 Digital Sun Sensor 2 0.06 0.12 Star Tracker 1 0.18 0.18 Coarse Three Axis Gyro 1 0.11 0.11 3 axis Magnetometer 1 0.05 0.05 GNSS 1 0.18 0.18 Three Reaction Wheel Package 1 0.46 0.46 Three Axis Magnetorquer 1 0.10 0.10 Power 1.75 Battery 1 0.27 0.27 Solar Array 5 0.18 0.88 PCDU 5 0.12 0.60 Table 5-1: Module selection – Scenario 1 (LEO w/o propulsion) 5.5.1.3 Configuration An exploded and closed view of the spacecraft are presented below (note the undefined payload volume is represented here as a black box):

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Figure 5-1: Configuration – Scenario 1 (LEO w/o propulsion) 5.5.1.4 Performance • Available payload volume: 22 L (230 x 280 x 340 mm) • Available payload mass: 10.88 kg (assuming total S/C launch mass of 20 kg) • Available payload power: 10 W (nominal & eclipse, with 50% duty cycle) • Communications: For a link budget at 800 km altitude with 30 min of contact per day at Redu, the amount of downloaded data is 4.5 Gb per day. • AOCS performance:

AOCS Performance APE [arcmin] RPE [arcsec]/100s Determination 0.4 4 Pointing 1 10 Slewing 90 deg in < 100 sec

Table 5-2: AOCS performance – Scenario 1 (LEO w/o propulsion) 5.5.1.5 Mass Budget Table 5-3 defines the mass budget for the Scenario 1 design. As can be seen, the spacecraft is well within the limit of 20 kg, allowing a payload mass of up to 10.9 kg to stay within the original mass envelope.

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NanoSat

Target Spacecraft Mass at Launch 20.00 kg Below Mass Target by: 10.88 kg

Without Margin Margin Total % of Total Dry mass contributions %kgkg Structure 2.27 kg 0.00 0.00 2.27 29.86 Thermal Control 0.39 kg 0.00 0.00 0.39 5.13 Mechanisms 0.06 kg 0.00 0.00 0.06 0.80 Communications 0.99 kg 0.00 0.00 0.99 13.02 Data Handling 0.95 kg 0.00 0.00 0.95 12.49 AOCS 1.19 kg 0.00 0.00 1.19 15.67 Propulsion 0.00 kg 0.00 0.00 0.00 0.00 Power 1.75 kg 0.00 0.00 1.75 23.02 Harness 0.00 kg - - - - Total Dry(excl.adapter) 7.60 7.60 kg System margin (excl.adapter) 20.00 %1.52kg Total Dry with margin (excl.adapter) 9.12 kg Other contributions Wet mass contributions Propellant 0.00 kg 0.00 0.00 0.00 0.00 Adapter mass (including sep. mech.), kg 0.00 kg 0.00 0.00 0.00 0.00 Total wet mass (excl.adapter) 9.12 kg Table 5-3: Mass budget – Scenario 1 (LEO w/o propulsion) Note from the above table that, while a 20% systems level margin has been applied to the spacecraft dry mass, no margins have been applied at subsystem/unit level. This is in keeping with the design philosophy as identified in Section 5.2 above. A diagrammatic representation of the system dry mass is provided in Figure 5-2 below:

Power 23% Structure 30%

AOCS Thermal Control 16% 5%

Me c h a n i s m s 1% Data Handling 12% Communications 13% Figure 5-2: Scenario 1 dry mass breakdown 5.6 Mission Baselines - Scenario 2 (LEO w/ Propulsion) The second test case was similar to Scenario 1, albeit with onboard propulsion included. The scenario definition is: • LEO 600 km altitude circular SSO (10:30 LTAN) o (Link Budget to consider 800 km altitude) • Onboard propulsion for 1 DoF control

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o (max. 5 kg total propulsion S/S mass allocated) • 3-Axes Stabilised - Nadir Pointing Spacecraft • P/L Input Power: 5 W (continuous) or 10 W (50% Duty Cycle) • 2 Mbps Data rate. 5.6.1 System Baseline Design 5.6.1.1 Description Taking into account the mission objectives, requirements and assumptions outlined above, the following system baseline was defined. 5.6.1.2 Module Selection / Equipment List The following list presents the module selection for the NanoSat spacecraft in this scenario. Note the module selection justifications are given in the individual subsystem sections, and are not repeated here.

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Element 1 - NanoSat FUNCTIONAL SUBSYSTEM nr Mass (kg) per unit Total Mass (kg) Structure 2.27 Side Panel 4 0.16 0.62 Panel Brackets 8 0.03 0.24 Top Panel 1 0.10 0.10 Harness 1 0.60 0.60 Half Panels 4 0.08 0.31 Bottom Panel 1 0.15 0.15 Horizontal Brackets - Long 8 0.02 0.16 Horizontal Brackets - Short 8 0.01 0.08 Thermal Control 0.39 Black paint 1 0.30 0.30 Heater line (2 heaters+1 sensor) 1 0.09 0.09 Mechanisms 0.06 Nano-Terminator Deorbit Module (NTDM) ( incl nan 1 0.06 0.06 Communications 0.99 Light S-band LGA 2 0.02 0.04 Cables 1 0.05 0.05 UP/DW converter 2W 1 0.90 0.90 Data Handling 0.95 DCU 1 0.70 0.70 Spacewire interface 1 0.05 0.05 TB interface 1 0.10 0.10 RF electronics 1 0.10 0.10 AOCS 1.19 Digital Sun Sensor 2 0.06 0.12 Star Tracker 1 0.18 0.18 Coarse Three Axis Gyro 1 0.11 0.11 3 axis Magnetometer 1 0.05 0.05 GNSS 1 0.18 0.18 Three Reaction Wheel Package 1 0.46 0.46 Three Axis Magnetorquer 1 0.10 0.10 Propulsion 1.81 Propellant Tank Module (A) Tank 1 1.25 1.25 Propellant Tank Module (A) Fill/Vent Valve 1 0.07 0.07 Propellant Tank Module (A) Fill/Drain Valve 1 0.07 0.07 Propellant Tank Module (A) Filter 1 0.10 0.10 Propellant Tank Module (A) Helium/Pressurant 1 0.00 0.00 1 Thruster Butane Module (E) Thruster 1 0.08 0.08 1 Thruster Butane Module (E) Vaporiser 1 0.05 0.05 1 Thruster Butane Module (E) Electronics 1 0.10 0.10 1 Thruster Butane Module (E) Piping 1 0.05 0.05 1 Thruster Butane Module (E) Wiring 1 0.05 0.05 Power 1.75 Battery 1 0.27 0.27 Solar Array 5 0.18 0.88 PCDU 5 0.12 0.60 Propellant 1.99 Table 5-4: Module selection – Scenario 2 (LEO w/ propulsion)

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5.6.1.3 Configuration The exploded and closed views of the spacecraft are presented below (note the undefined payload volume is represented as a black box). The payload here is assumed to be 100% contained within the S/C volume but, as the top panel is not used for unit accommodation other than payload mounting, it could, in principle, extend beyond this. No further assessment of this has been done in this study:

Figure 5-3: Configuration – Scenario 2 (LEO w/ propulsion) 5.6.1.4 Performance • Available payload volume: 8 L (250 x 280 x 120 mm) • Available payload mass: 6.71 kg (assuming total S/C launch mass of 20 kg) • Available payload power: • Available ∆V for 1 DoF orbit control (using Cold Gas Butane, Isp = 62 s): 62.50 m/s • Communications: • AOCS performance: 5.6.1.5 Mass Budget Table 5-5 defines the mass budget for the Scenario 2 design. As can be seen, the spacecraft is well within the limit of 20 kg, allowing a payload mass of up to 6.7 kg to stay within the original mass envelope.

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NanoSat

Target Spacecraft Mass at Launch 20.00 kg Below Mass Target by: 6.71 kg

Without Margin Margin Total % of Total Dry mass contributions %kgkg Structure 2.27 kg 0.00 0.00 2.27 24.11 Thermal Control 0.39 kg 0.00 0.00 0.39 4.14 Mechanisms 0.06 kg 0.00 0.00 0.06 0.65 Communications 0.99 kg 0.00 0.00 0.99 10.52 Data Handling 0.95 kg 0.00 0.00 0.95 10.09 AOCS 1.19 kg 0.00 0.00 1.19 12.66 Propulsion 1.81 kg 0.00 0.00 1.81 19.25 Power 1.75 kg 0.00 0.00 1.75 18.59 Harness 0.00 kg - - - - Total Dry(excl.adapter) 9.42 9.42 kg System margin (excl.adapter) 20.00 %1.88kg Total Dry with margin (excl.adapter) 11.30 kg Other contributions Wet mass contributions Propellant 1.99 kg 0.00 0.00 1.99 14.99 Adapter mass (including sep. mech.), kg 0.00 kg 0.00 0.00 0.00 0.00 Total wet mass (excl.adapter) 13.29 kg Table 5-5: Mass budget – Scenario 2 (LEO w/ propulsion) Note, margin policy is as for Scenario 1. A diagrammatic representation of the system dry mass is provided in Figure 5-4 below: Power 19% Structure 23%

Thermal Control Propulsion 4% 19% Mechanisms 1%

Communications 11%

AOCS Data Handling 13% 10%

Figure 5-4: Scenario 2 dry mass breakdown 5.7 Mission Baselines - Scenario 3 (GTO) The final test case represented a GTO scenario, with an inertially-pointing black-box, undefined payload. De-orbiting within 25 years is considered an essential mission requirement. The scenario definition is: • Injection orbit: 250 x 35143 km o with 6º inclination wrt Equator • Final orbit (example): 600 x 35143 km o 10.4 hrs period (max eclipse duration 2.2 hrs)

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• Onboard propulsion for 3 DoF control and momentum offloading o (max. 5 kg wet mass allocated) • 3-Axes Stabilised – Inertial/Sun Pointing Spacecraft (Observatory type application) • P/L Input Power: 10 W continuous • 2 Mbps Datarate. 5.7.1 System Baseline Design 5.7.1.1 Description Taking into account the mission objectives, requirements and assumptions outlined above, the following system baseline was defined. 5.7.1.2 Module Selection / Equipment List The following list presents the module selection for the NanoSat spacecraft in this scenario. Note that the module selection justifications are given in the individual subsystem sections, and are not repeated here.

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Element 1 - NanoSat FUNCTIONAL SUBSYSTEM nr Mass (kg) per unit Total Mass (kg) Structure 2.69 Side Panel 4 0.16 0.62 Panel Brackets- vertical 8 0.03 0.24 Top Panel 1 0.15 0.15 Harness 1 0.60 0.60 Half Panels 4 0.08 0.31 Bottom Panel 1 0.21 0.21 Deployable Panel 2 0.16 0.31 Panel Brackets- horizontal long 8 0.02 0.16 Panel Brackets- horizontal short 8 0.01 0.08 Thermal Control 1.37 Black paint 1 0.30 0.30 Heater line (2 heaters+1 sensor) 1 0.09 0.09 Heat pipe 3 0.20 0.60 MLI blankets (mass is per m2) 1 0.50 0.38 Mechanisms 0.33 S/A Deployment Mechanism SDM (w/o HDRM a 2 0.09 0.17 Hold-Down and Deployment Mechanism (HDRM 4 0.04 0.16 Communications 0.99 UP/DW converter 4WRF 1 0.90 0.90 Lightweight S-band LGA 2 0.02 0.04 Cables 1 0.05 0.05 Data Handling 0.95 DCU 1 0.70 0.70 Spacewire interface 1 0.05 0.05 TB interface 1 0.10 0.10 RF electronics 1 0.10 0.10 AOCS 0.96 Digital Sun Sensor 2 0.06 0.12 Star Tracker 1 0.18 0.18 Coarse Three Axis Gyro 1 0.11 0.11 Three Reaction Wheel Package 1 0.46 0.46 Three Axis Magnetorquer 1 0.10 0.10 Propulsion 3.21 Propellant Tank Module (A) Tank 1 1.25 1.25 Propellant Tank Module (A) Fill/Vent Valve 1 0.07 0.07 Propellant Tank Module (A) Fill/Drain Valve 1 0.07 0.07 Propellant Tank Module (A) Filter 1 0.10 0.10 Propellant Tank Module (A) Helium/Pressurant 1 0.00 0.00 Solid Rocket Module (C) Motor 1 0.48 0.48 Solid Rocket Module (C) Nozzle 1 0.10 0.10 Solid Rocket Module (C) Ignition/Electronics 1 0.10 0.10 1 Thruster Butane Module (E) Thruster 1 0.08 0.08 1 Thruster Butane Module (E) Vaporiser 1 0.05 0.05 1 Thruster Butane Module (E) Electronics 1 0.10 0.10 1 Thruster Butane Module (E) Piping 1 0.05 0.05 1 Thruster Butane Module (E) Wiring 1 0.05 0.05 2 Thruster Butane Module (F) Thruster 1 0.23 0.23 2 Thruster Butane Module (F) Vaporiser 1 0.05 0.05

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Element 1 - NanoSat FUNCTIONAL SUBSYSTEM nr Mass (kg) per unit Total Mass (kg) Propulsion 3.21 2 Thruster Butane Module (F) Electronics 1 0.15 0.15 2 Thruster Butane Module (F) Piping 1 0.15 0.15 2 Thruster Butane Module (F) Wiring 1 0.15 0.15 Power 2.51 Battery 6 0.27 1.62 Solar Array 3 0.18 0.53 PCDU 3 0.12 0.36 Propellant 1.99 Table 5-6: Module selection – Scenario 3 (GTO) 5.7.1.3 Configuration Internal and external views of the spacecraft are presented below (note the undefined payload volume is represented as a black box) in Figure 5-5. Note that this configuration involves two fixed-deployable solar arrays to accommodate the increased power demand of the GTO scenario.

Figure 5-5: Configuration – Scenario 3 (GTO – internal payload) Furthermore, as commented in the Configuration chapter, it is possible to externally mount the payload on the top face of the platform. This gives options such as the one identified in Figure 5-6 below:

Figure 5-6: Configuration – Scenario 3 (GTO – external payload) 5.7.1.4 Performance • Available payload volume: 8 L (250 x 280 x 120 mm)

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• Available payload mass: 1.5 kg (approx) • Available payload power: 10 W (continuous) • Communications: Communications downlink uses Variable Coding and Modulation (VCM). Considering a link budget for Malindi with a transmission window between 3000 and 11000 km (37 min/day), the use of VCM gives an amount of downloaded data of 1.1 Gb per day. • Available ∆V: o 3 DoF control (using Cold Gas Butane, Isp = 62 s): ∆V = 62.5 m/s o 1 DoF solid propellant de-orbiting module (Isp = 260 s): ∆V = 50 m/s • AOCS performance:

AOCS Performance APE [arcmin] RPE [arcsec]/100s Determination 0.4 4 Pointing 1 10 Slewing 90 deg in < 100 sec

Table 5-7: AOCS performance – Scenario 3 (GTO) 5.7.1.5 Mass Budget Table 5-8 defines the mass budget for the Scenario 3 design. As can be seen, the spacecraft is well within the limit of 20 kg allowed. NanoSat

Target Spacecraft Mass at Launch 20.00 kg Below Mass Target by: 1.53 kg

Without Margin Margin Total % of Total Dry mass contributions %kgkg Structure 2.69 kg 0.00 0.00 2.69 19.57 Thermal Control 1.37 kg 0.00 0.00 1.37 9.94 Mechanisms 0.33 kg 0.00 0.00 0.33 2.42 Communications 0.99 kg 0.00 0.00 0.99 7.21 Data Handling 0.95 kg 0.00 0.00 0.95 6.92 AOCS 0.96 kg 0.00 0.00 0.96 7.02 Propulsion 3.94 kg 0.00 0.00 3.94 28.67 Power 2.51 kg 0.00 0.00 2.51 18.26 Harness 0.00 kg - - - - Total Dry(excl.adapter) 13.74 13.74 kg System margin (excl.adapter) 20.00 %2.75kg Total Dry with margin (excl.adapter) 16.48 kg Other contributions Wet mass contributions Propellant 1.99 kg 0.00 0.00 1.99 10.78 Adapter mass (including sep. mech.), kg 0.00 kg 0.00 0.00 0.00 0.00 Total wet mass (excl.adapter) 18.47 kg Launch mass (including adapter) 18.47 kg Table 5-8: Mass budget – Scenario 3 (GTO) Note, margin policy is as for Scenario 1. A diagrammatic representation of the system dry mass is provided in Figure 5-7 below:

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Power Structure 19% 21%

Thermal Control 10%

Propulsion Me c h a n i s m s 25% 3%

Communications 8% AOCS Data Handling 7% 7% Figure 5-7: Scenario 3 dry mass breakdown

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6 CONFIGURATION 6.1 Requirements and Design Drivers The requirements and design drivers for the configuration are: • The spacecraft shall be as small as possible • The configuration shall be modular, with 100% recurrent design of modules • The configuration shall accommodate all subsystem equipment and instruments • The configuration shall provide unobstructed field of view for instruments, sensors and antennas, and allow free deployment of mechanisms • The configuration shall take into account structural considerations • The configuration shall take into account AIT/AIV considerations. 6.2 Assumptions and Trade-Offs This study is to investigate the smallest practical flexible platform, with a “lego style” configuration and assembly approach. The payload design is outside the frame of this study and is assumed as a black box with a density of 1 kg/dm3. The design of the separation system of the spacecraft from the launcher is outside the frame of this study. It is nevertheless assumed a 3-point bolt release interface (from the type of the SAAB SS175) is to be used and located on the base plate of the NanoSat. Alternatives (e.g. rail systems) should not be discounted. 6.3 Baseline Design The baseline configuration of the spacecraft was chosen as a box with bevel edges resulting in an octagonal sided box as shown in (Figure 6-1). The following considerations have been taken into the selection process: • Curved surfaces (sphere, cylinder) present complicated shapes for manufacturing and solar cell mounting and should therefore be discarded • The platform shall be suitable for different mission types and exotic, dedicated shapes should be avoided • There should be mounting possibilities for equipment with free views on every side of the spacecraft • The design shall be structurally favourable (simple and direct load path, compact shape, symmetric).

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Figure 6-1: Baseline spacecraft

The design driver for the side panel size is the required solar cell area to meet the power demands. One solar array panel has the min. dimensions (w) 230 mm × (h) 340 mm. For attachment to neighbouring panels the SA panel size is (w) 230 mm × (h) 370 mm. The different elements required for the entire baseline configuration, excluding interfaces, are reduced to four: • Side panel – fully recurring module • Half side panel – fully recurring module • Top panel – bespoke per mission • Base panel – bespoke per mission. The half side panels provide the mounting area for equipment that requires a specific pointing in the given direction. These panels provide attachment means on the inner and on the outer surface. For cases in which an instrument (AOCS sensor or payload) is to protrude the panel, predefined cut-outs may be implemented that are an integral part of the structure but can easily be removed if necessary. Predefined attachment and cut-outs locations are also indicated on the top and bottom panels for optional propellant tubing, sensors or antennas. A predefined standard equipment mounting hole pattern is forseen, allowing standard, pre-set mounting points/inserts to be incorporated into the panel.

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6.4 Scenario Study Case 1 - LEO (w/o Propulsion) 6.4.1 Design Drivers The design drivers for the LEO case were: • A low centre of mass • Maximization of the available payload envelope • Functional accommodation of the subsystems elements. 6.4.2 Description In this case the payload has an optical opening through a half-side panel to provide nadir- pointing viewing capabilities, as an example for an earth observation mission. The Nadir facing half-panel also accommodates the Nano-Deorbit module and one LGA (Figure 6-2). For Omni- directional coverage the 2 LGA’s are placed on opposite sides of the spacecraft. The subsystem elements are distributed and attached on the back of the side panels by means of the panel inserts (Figure 6-4). Few elements have been accommodated on the half-panels when a specific FoV is required e.g.: Sun sensor, Star tracker and LGA (Figure 6-2 & Figure 6-3).

Star tracker

Payload optics

Nano-deorbit module

LGA

Figure 6-2: LEO (Noon-Midnight) from Nadir side

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Sun sensors

LGA

Figure 6-3: LEO (Noon-Midnight) from Zenith side

DCU UP/DW converter 2W Star tracker

PCU MPPT(s) Magnetorquers

GNSS, Gyro, Magnetometer

Battery Reaction wheels

Figure 6-4: LEO (Noon-Midnight) inner configuration

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Figure 6-5: LEO (Dawn-Dusk) from Nadir side 6.4.3 Modules Selection (With Justification) The modules that have been implemented in the LEO study case are defined by each of the spacecraft subsystems. Figure 6-4 above shows the accommodation of the elements in the spacecraft. 6.4.4 Achievable Performances/Realised Functionalities For the LEO case 1, the envelope available for the payload was derived as: (l) 280mm × (w) 230mm × (h) 340mm with a volume of 0.022 m3 (22 Litres). 6.5 Scenario Study Case 2 - LEO (w/ Propulsion) 6.5.1 Design Drivers The design drivers are as for case 1, except the platform must also now accommodate a propulsion system. 6.5.2 Description The tank is polar mounted in horizontal position on the base plate together with most of the propulsion equipment (see Figure 6-8). The image below (Figure 6-6) shows a 1 DoF thruster accommodation for pure translation.

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Fill/Vent & Fill/Drain Valves (2)

Thruster flight direction

Figure 6-6: LEO case 2 - 1 DoF thruster accommodation Optionally, the thruster accommodation shown in Figure 6-7 allows optimized torque around each axis of the spacecraft plus torque free translation in one direction.

Thrusters Roll

Fill/Vent & Fill/Drain Valves (2) Thrusters Pitch

Roll Thrusters Yaw

Pitch

Thruster flight

Yaw direction Figure 6-7: LEO case 2 - 3 DoF thruster accommodation (Optional)

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Propulsion tank Propulsion electronics

Figure 6-8: LEO case 2 accommodation (with propulsion module) 6.5.3 Modules Selection (With Justification) The modules that have been implemented in the LEO study case are defined by each of the spacecraft subsystems. The Figure 6-8 above shows the accommodation of the elements in the spacecraft. 6.5.4 Achievable Performances/Realised Functionalities For the LEO case with propulsion the dimensional envelope available for the payload was derived as: (l) 280mm × (w) 250mm × (h) 120mm with a volume of 0.008 m3 (8 Litres). The study case includes one thruster for translational force. Alternatively, it can be equipped with a three axis attitude control propulsion system for orbital- manoeuvres, station-keeping and wheel-offloading. 6.6 Scenario Study Case GTO 6.6.1 Design Drivers The design drivers for the GTO case were: • A low centre of mass • Maximization of the available payload envelope • Functional accommodation of the subsystems elements • Inclusion of two deployable SA panels

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• Account for a solid propellant deorbit propulsion module. 6.6.2 Description In this case the payload has been assumed with sensors on the top panel. Similar to the LEO case, omni-directional coverage is achieved with 2 LGA’s placed on opposite sides of the spacecraft. The subsystems elements are distributed and attached on the back of the side panels by means of the panel inserts (Figure 6-8). Few elements have been accommodated on the half- panels when a specific FoV is required e.g.: Sun sensors & Star tracker.

SDM (4)

Figure 6-9: GTO case with deployed SA panels and internal payload

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Figure 6-10: GTO case with deployed SA panels and external payload (optional)

Solid propellant deorbit module

Figure 6-11: GTO location of solid propulsion deorbit module

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LGA

PW/DW Batteries converter 4WRF

Star tracker

PCU Reaction MPPT wheels, Magnetorquer

SA Solid HDRM propellant (4) deorbit

module Figure 6-12: GTO inner accommodation The mechanisms for the deployment of the Solar Array panels (SDM and HDRM) are mounted on the wall panels lying directly underneath. 6.6.3 Modules Selection (With Justification) The modules that have been implemented in the GTO study case are defined by each of the spacecraft subsystems. The Figure 6-12 above shows the accommodation of the subsystem units in the spacecraft. 6.6.4 Achievable Performances/Realised Functionalities For the GTO study case the dimensional envelope for a payload inside of the spacecraft was derived as: (l) 280mm × (w) 250mm × (h) 120mm with a volume of 0.008 m3 (8 Litres). The present configuration also allows for a payload (or payload sensors) that is placed on the exterior (e.g. on the top panel - provided that this is not used for solar arrays). A solid propellant deorbit module has also been considered and preliminarily positioned to achieve alignment of the CoM and the thrust vector. Furthermore two deployable solar panels have been implemented. 6.7 Overall Dimensions Figure 6-13 and Figure 6-14 show the overall dimensions of the LEO case with propulsion.

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Figure 6-13: LEO extended side view

Figure 6-14: LEO extended top view Figure 6-15 and Figure 6-16 show the overall dimensions of the GTO case with two deployable SA panel.

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Figure 6-15: GTO deployed configuration side view

Figure 6-16: GTO stowed configuration top view

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6.8 Options The components used in the baseline configuration can also be recombined into other shapes shown below (Figure 6-17).

Figure 6-17: Configuration options The options in the second row are exclusively made of the four base elements that constitute the baseline model and do not require any customized elements. The options in the first row will require an additional custom panel. 6.9 Technology Requirements There are no technological requirements for the configuration subsystem.

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7 STRUCTURES 7.1 Requirements and Design Drivers • Provide mechanical support and containment for all spacecraft units, equipment and subsystems (Interfaces with all subsystems) • Modular approach o Use of recurrent elements o Standard subsystem attachment points • Smallest possible design envelope o Miniaturization study • Consider structural issues o Location of CoM, load paths,… o Size for 20 kg with 4 panels • Allow free movement of mechanisms and deployables. 7.2 Sub-System Description The structure subsystem encompasses all of the structural aspects of the platform. To achieve modularity, standardised panels have been designed which will be mass producible. There is a mass disadvantage to this approach as each panel will be designed for the worst case but this is to be expected with a modular reproducible design. The structure can be broken down into several modules, including panels and panel interfaces. Specifically the modules anticipated are: • Wall Panels • Half Wall Panels • Top Panels • Bottom Panels • Panel Interfaces. To achieve modularity and reproducibility, each panel will contain interfaces that will match with any mission scenario. The only variation will be in the configuration of the spacecraft, which will determine, how many of each panel will be required. In this sense it may be necessary to produce Top and Bottom Panels tailored to the mission. Currently the configuration is an octagonal structure, but with the modularity of the side panels, it is completely feasible, to adopt a cubic or hexagonal structure. This could be achieved without any changes to the production of the Wall Panels, Half Wall Panels and the Panel Interfaces. The purpose of each module is indicated in Table 7-1. Not all functions are used for every mission scenario.

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Module Purpose Wall Panel Primary Load Path, Solar Cell Mounting Surface, Unit and Equipment Mounting Surface, Deployable Solar Panel Half Wall Panel Closure Panels for Octagonal Spacecraft, Thruster Mounting Surface, Solar Array Hold-down Surface*, Can contain cutouts for Instrument Field of View Top Panel Closure Panel, Interface with Payload, Solar Cell Mounting Surface Bottom Panel Launcher Interface, Interface with Propulsion System, Interface with Payload Panel Interfaces Connection between all Panels Table 7-1: Purpose of Modules * The position of the Solar Array hold downs would be the subject of detailed design to determine the best location. It is considered best to locate all hold downs and hinges on the same panel, whether that panel is the Wall or Half Wall Panel. 7.2.1 Interfaces The Structure subsystem interfaces with all of the other subsystems. Within the scope of this study, most interfaces have been considered, however the Launcher interface was not investigated in any detail. Four standard interface points are indicated on the Bottom Panel, but no candidate launchers or launch adaptors have been investigated. 7.2.1.1 Wall Panels The design of the structure panels will allow the integration of all subsystems to standard interface patterns on the inside of the spacecraft. On the external face of the panels solar cells will be mounted. In the mission scenario when deployable panels are required, identical panels can be used. For the fixed wall panel, no solar cells will be attached on the external face, while for the deployed panel, no equipments shall be mounted on the rear face. The deployable solar array can have attachment points on either the vertical edge or the horizontal edge depending on the needs of the mission.

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Figure 7-1: Wall Panel with a standard Interface Pattern

Figure 7-2: Wall Panel interface to Solar Cells 7.2.2 Half Wall Panels The design of the half wall panels allows for cut-outs for Payload Field of View. If the cut-outs are not used, the panels can mount smaller equipments for other subsystems, such as thrusters, antennas and de-orbit mechanisms. It can be considered that these panels would also provide the

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 72 of 267 hold-down and deployment mechanisms for any deployable arrays, however, it may be more mass effective to contain all of these mechanisms on the main wall panels.

Figure 7-3: Half Wall Panel with potential Cut-out Locations 7.2.3 Top Panel The design of the Top panel allows for Payload mounting and also contains four cut outs for Antennas. The modular approach provides standard interface points for the payload and four standard interface cut out locations for the antennas as can be seen below in Figure 7-4: . For some mission scenarios, the Payload would be mounted internally, thus the Payload mounting points would not be needed. This would allow for the integration of solar cells on the top panel as seen below in Figure 7-5: . Despite the different functions for the Top Panel for different mission scenarios, the Panel is mass producible. The concept is to have locations and inserts for all possible cut-outs, but only to have a cut-out when necessary, thereby allowing the placement of solar cells.

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Figure 7-4: Top Panel with Payload Interfaces and Cut Out Locations (GTO Mission)

Figure 7-5: Top Panel for a LEO Mission 7.2.4 Bottom Panel The design of the Bottom panel allows for Launcher Interface, Payload mounting and also contains two cut outs for Antennas. The modular approach provides standard interface points for the payload and four standard interface cut out locations for the antennas as can be seen below in Figure 7-6: . For some mission scenarios, the Payload would be mounted on the Top Panel, thus the Payload mounting points would not be needed. Despite the different functions for the Bottom Panel for different mission scenarios, the Panel is mass producible. The concept is to have locations and inserts for all possible Cut-outs, but only to have a cut-out when necessary.

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Figure 7-6: Bottom Panel Interfaces 7.2.5 Panel Brackets The concept for the connection of the panels can either be a conventional approach or an innovative approach. For the conventional approach brackets would be utilised. A standard bracket shape would be defined and then depending on the location, different lengths would be used. For the spacecraft configuration baselined in this study, those lengths would be ~115, 230 and 350 mm. A conventional bracket as seen below in Figure 7-7: could be used.

Figure 7-7: Conventional Bracket For an innovative approach, integrating the connection into the panel structure would reduce the number of parts and improve the mass efficiency. The possible concepts would have to be investigated in detail to arrive at the best solution. Some concepts are included below.

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Figure 7-8: Integrated Panel Connection Concept

Figure 7-9: Detail of toothed connection The concept with connection shown above is to allow for any configuration. The angle between the two panels can vary depending on the configuration, so this concept allows for 45º, 60º and 90º angles (and multiples thereof) between the panels. A toothed rod would be inserted for fixation.

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7.3 Module Descriptions 7.3.1 Module 1 (Wall Panel) 7.3.1.1 Description and Overview The design of a conventional Wall Panel would be a simple task with little development costs, high TRL and low risk. A standard honeycomb panel, with CFRP or Aluminium facesheets could be easily manufactured in bulk. However, within the time frame of 8-10 years, it would be advantageous to investigate innovative possibilities. Alternative materials and manufacturing methods could be investigated. Areas for development include the integration of additional functionality within the structure. Conventionally spacecraft structures are designed to withstand the most severe loading situation, usually the launch, but spend the remainder of the mission oversized. Embedding extra functionality into the structure, whether in the form of embedded batteries or integrated harness, would lead to a more efficient structure, with mass saving benefits. On the wall panel, which will be generic, there will have to be the possibility to integrate the hinge and the release mechanism on the inner side of the panel for deployability. This however, should not interfere with the use of the wall panel in a non-deployed configuration. There are four considerations for the use of deployable panels: • The hold down on the fixed panel • The hinge on the fixed panel • The hinge on the deployed panel • The release on the deployed panel. A trade off should be performed to determine the location of the hold downs and hinges for the deployable panels. They can either be located on the wall panels or the half wall panel, or even a combination of both, for example, the hinge on the wall panel and the hold down on the half wall panel. The mass implications for each configuration should be investigated, both in the scenario where the deployment is needed and in the scenario where it is not needed. 7.3.1.2 Key characteristics • Mass: 0.156 kg • Power: Not applicable • Dimensions: 230 X 340 X 10 mm. 7.3.1.3 Scalability This module cannot be scaled. 7.3.1.4 Multi mission suitability This module can be used in any mission scenario.

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7.3.2 Module 2 (Half Wall Panel) 7.3.2.1 Description and Overview The half wall panel concept is identical to the wall panel concept. The only difference is how the panels will be loaded. The Wall Panels will carry the majority of the subsystem equipments, while the half wall panels will have optional cutouts and fixing locations for smaller items such as thrusters. From a material perspective only one study would need to be performed to determine the best solution for all panel types. 7.3.2.2 Key Characteristics • Mass: 0.078 kg • Power: Not applicable • Dimensions: 115 X 340 X 10 mm. 7.3.2.3 Scalability This module could be scaled (doubled) to replace Module 1 (Wall Panel). Adaption would be needed to connect the two half panels. 7.3.2.4 Multi mission suitability This module can be used on any mission scenario. 7.3.3 Module 3 (Top Panel) 7.3.3.1 Description and Overview The top panel design will be more dependent on the mission than the wall panels. In the LEO Scenario, the top panel would be populated with Solar Cells, while for the GTO Scenario the top panel would have Payload cutout locations. This could be incorporated in a single panel design. This would not be very mass effective as the panel that is capable of supporting the payload would be approximately 50% heavier than a panel which is lightly loaded. The development work would be based on the concept rather than the configuration so a single study process would incorporate all panel modules. From a material perspective only one study would need to be performed to determine the best solution for all panel types. 7.3.3.2 Key Characteristics • Mass Range: 0.10 – 0.15 kg • Power: Not applicable • Dimensions: Depends on the configuration chosen. 7.3.3.3 Scalability This module cannot be scaled. 7.3.3.4 Multi mission suitability This module can be used on any mission scenario.

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7.3.4 Module 4 (Bottom Panel) 7.3.4.1 Description and Overview The bottom panel design will be more dependent on the mission than the wall panels. In the LEO Scenario, the bottom panel would have as its primary function provision of the launcher interface and the Payload support, while for the GTO Scenario the bottom panel would have a propulsion system mounted rather than the Payload interface. This could be incorporated in a single panel design. However a family of panels which are tailored to a few specific functions, would provide mass efficiency but also the repeatability for low cost large scale production. The development work would be based on the concept rather than the configuration so a single study process would incorporate all panel modules. From a material perspective only one study would need to be performed to determine the best solution for all panel types. 7.3.4.2 Key Characteristics • Mass: 0.15 – 0.21 kg • Power: Not applicable • Dimensions: Depends on the configuration chosen. 7.3.4.3 Scalability This module cannot be scaled. 7.3.4.4 Multi mission suitability This module can be used on any mission scenario. 7.3.5 Module 5 (Panel Interfaces) 7.3.5.1 Description and Overview The panel interface design could be a conventional option such as an angle bracket. This would be easily mass producible and produced in long batches and just be cut to length as required. This would not be the most mass effective method as inserts and interfaces would be needed on each panel edge to connect to the brackets. To aid with the mass production of the panels, the panel inserts would have a standard layout which would allow the attachment of any unit. The requirement on the unit supplier would be to have a standard interface. The concept is illustrated in Figure 7-1. The main use would be for the wall and half wall panels, as the top and bottom panels may be mission dependent. A family of panels with standard inserts patterns for specific missions, could be achieved. The inserts would be integrated within the wall (&half wall) panels irrespective of whether units were to be attached or not. If units are to be attached then the holes can be drilled out for the interface, if not, as for example in the case where the wall panels are to be used as solar panels, then the panel would not need to be drilled. Another concept as shown above in Figure 7-8, is to incorporate the panel interfaces into the panel design.

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7.3.5.2 Key Characteristics • Mass: 0.01 – 0.02 kg for conventional bracket, depending on length • Power: Not applicable • Dimensions: Depends on the configuration chosen. 7.3.5.3 Scalability This module cannot be scaled. 7.3.5.4 Multi mission suitability This module can be used on any mission scenario. 7.3.6 Module 6 (Harness) The concept for the harness is to minimise mass and improve efficiency. Space Wire will only be used for “Intelligent” connections as it is very heavy per unit length. Using conventional methods, the harness may be longer due to inefficiencies of connection. An improvement on this concept would be to use a serial bus which would provide a standard harness connection on every panel. Rather than consider the harness once the mission has been defined and the necessary units determined, by using a serial bus, any combination of units could be connected to power and data-handling as needed. Already at a reasonably high TRL level, micro connectors provide mass savings over conventional connectors. The current TRL of these devices is 6. However, due to the size of the NanoSat spacecraft and the mass minimisation effort, an innovative approach shall be to investigate the use of nano-tubes which could introduce significant mass savings. This has a high risk due to the very low TRL currently. As yet this technology is unproven but within the timeframe of this program, it should be possible to prove this technology – or eliminate it as a potential option. Nano-D Connectors would also introduce significant mass savings even with respect to micro connectors and these are currently at TRL 4 and expected to be TRL 6 within 2 to 3 years. 7.4 Scenario Study Case 1 (LEO w/o Propulsion) 7.4.1 Design Drivers Support of the payload and other subsystems. 7.4.2 Module Selection (Justification) All modules are selected. Specifically: • 4 x Wall Panels • 4 x Half Wall Panels • 1 x Top Panel • 1 x Bottom Panel • 24 x Conventional Brackets (0 otherwise if integrated into Panels).

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7.4.3 Mass Budget

Module nr Mass (kg) per unit Total Mass (kg) Margin (%) Margin (kg) Mass (kg) with Margin Side Panel 4 0.16 0.62 0.00 0.00 0.62 Panel Brackets 8 0.03 0.24 0.00 0.00 0.24 Top Panel 1 0.10 0.10 0.00 0.00 0.10 Harness 1 0.60 0.60 0.00 0.00 0.60 Half Panels 4 0.08 0.31 0.00 0.00 0.31 Bottom Panel 1 0.15 0.15 0.00 0.00 0.15 Horizontal Brackets - Long 8 0.02 0.16 0.00 0.00 0.16 Horizontal Brackets - Short 8 0.01 0.08 0.00 0.00 0.08 Total 2.27 Table 7-2: Mass budget – Scenario 1 7.5 Scenario Study Case 2 (LEO w/ Propulsion) For the Structure subsystem, this case is assumed identical to Scenario Case 1 above.

7.6 Scenario Study Case 3 (GTO Scenario) 7.6.1 Design Drivers Support of the payload and other subsystems. 7.6.2 Module Selection (Justification) All modules are selected. Specifically • 6 x Wall Panels (incl. 2 deployable panels) • 4 x Half Wall Panels • 1 x Top Panel • 1 x Bottom Panel • 24 x Conventional Brackets (0 otherwise if integrated into Panels). 7.6.3 Mass/Power Budget

Module nr Mass (kg) per unit Total Mass (kg) Margin (%) Margin (kg) Mass (kg) with Margin Side Panel 4 0.16 0.62 0.00 0.00 0.62 Panel Brackets 8 0.03 0.24 0.00 0.00 0.24 Top Panel 1 0.15 0.15 0.00 0.00 0.15 Harness 1 0.60 0.60 0.00 0.00 0.60 Half Panels 4 0.08 0.31 0.00 0.00 0.31 Bottom Panel 1 0.21 0.21 0.00 0.00 0.21 Deployable Panel 2 0.16 0.31 0.00 0.00 0.31 Horizontal Brackets - Long 8 0.02 0.16 0.00 0.00 0.16 Horizontal Brackets - Short 8 0.01 0.08 0.00 0.00 0.08 Total 2.69 Table 7-3: Mass Budget – Scenario 3

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8 MECHANISMS 8.1 Requirements and Design Drivers The main design drivers for mechanisms to be used on NanoSat are, in order of precedence: 1. Lightness 2. Simplicity 3. Reliability There was no requirements specification identified for NanoSat, however the specific mechanisms designs will have to follow as a guideline the ESA ECSS E30-3 (Mechanisms design, see RD[4]). It has been defined that the separable interface mechanism to the launcher or dispenser is not to be considered as part of the mechanisms subsystem in this study. 8.2 Sub-System Description The following modules have been identified for the NanoSat mechanisms subsystem: • S/A Deployment Mechanism (SDM) • De-Orbit Deployment Mechanism (DDM) (optional if not using NTDM) • Hold-Down and Release Mechanism (HDRM) • Nano-Terminator De-Orbit Module (NTDM). It has to be noted that the DDM is only listed for completeness and to prove its existence, it is actually not used in this assessment, because it has been replaced by the NTDM. However it shall remain in this document and serve as backup. Table 8-1 provides a mapping and interdependency of the mechanisms to dedicated tasks:

Task on NanoSat Module to use Remark Deployment of Solar Array S/A Deployment Mechanism (SDM) Requires 2 HDRM’s per Panels SDM Hold-down and release of Hold-Down and Release Mechanism Option could be Nano- appendages (like S/A panel or (HDRM) release as used for NTDM DDM structure)

De-orbiting in LEO orbit Baseline: Nano-Terminator De-Orbit NTDM consists of Nano- Module (NTDM) terminator + Nano-release. Option requires 2 HDRM’s Option: De-Orbit Deployment per DDM Mechanism (DDM) Table 8-1: Mechanisms mapping to tasks 8.2.1 Interfaces The interfaces between the mechanism modules and the S/C are as follows:

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The following interfaces exist to other subsystems: • To Structure: SDM, DDM, HDRM, NTDM • To Power: HDRM, NTDM. The following interfaces exist between the mechanism modules: • HDRM to SDM and DDM. 8.3 Module Descriptions 8.3.1 Module 1: Solar Array Deployment Mechanism (SDM) 8.3.1.1 Description and Overview The task is a one-shot deployment of a single S/A panel, in order to optimize the sun incidence angle to get a better power output. The S/A and the HDRM is not part of this mechanism. Please refer to Figure 8-1 for a SDM concept overview. A system has been chosen which combines actuator and hinge function in one unit, and is realized using the so-called MAEVA hinge, see Figure 8-2. These items have already considerable flight heritage on (French) Microsats (e.g. MYRIADE platform). It is mandatory to accept that there is a “saloon-door”style deployment, i.e. there is an overshoot beyond the deployed position for up to 60 deg, and there will be up to 20 oscillations until steady-state deployed status. It is recommended to temporarily switch off AOCS during the deployment as the whole S/C might be moved, depending on the mass of the S/A panel. The main characteristics can therefore be summarized as follows: • Hinges only, 2 per SDM (no S/A and no HDRM) • Flexures, “saloon-door”-opening with oscillating over-swings • Hinge flexures for actuation and latch as well • The hinges might need to be further reduced in size to count for the step from Microsat to Nanosat, TBC • Hinge supplier: Metravib of France o http://www.01db-metravib.com/index.php?id=44&L=1 2 SDM’s shall be used for one S/C, each consisting of 2 hinges. The use of a speed regulator is not required due to the overswing nature (self dissipation of deployment energy) of the mechanism.

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Solar Array Deployment Mechanism (SDM) concept for NanoSat

Solar Array (not part of SDM, max. size assumed 34x23cm) Maeva-Hinge (2x)

Figure 8-1: SDM design concept

Figure 8-2: MAEVA-Hinge (from RD[5]) Note: French Patent 96.14416, Joint d’articulation automoteur, autoverrouillant et amortissant et articulation équipée de tels joints, European extension 97 947 079.6 and US extension 09/297,095. 8.3.1.2 Key characteristics The key characteristics are (extrapolated from existing device): • Mass: 86 g (for 2 hinges) • Dimensions: 0.3x0.02x0.02 m

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• Temp.: -75 to 105degC • Power: none. The main performance properties of one existing hinge are: • Angular position accuracy : < 1° • Driving torque : > 0.15 N.m • Open stiffness: 1000 N.m/rad • Holding torque : > 4.5 N.m

Figure 8-3: MAEVA-Hinge locking sequence (from RD[5]) 8.3.1.3 Scalability The SDM is scalable for different S/A panel sizes, or the number of hinges per panel can be adjusted. 8.3.1.4 Multi mission suitability There is no restriction for use on NanoSat missions. The fact of oscillating over-swing has to be accepted on system level (wrt. Collision risk with other appendages, or AOCS impacts).

8.3.2 Module 2: De-orbit Deployment Mechanism (DDM) 8.3.2.1 Description and Overview It has first to be noted that this mechanism for NanoSat is in fact superseded by the NTDM module. However a brief description is provided for completeness of the subsystem description. Please refer to Figure 8-4 for the DDM concept overview. The DDM uses the same hinges as the SDM, please refer to Figure 8-2 for more information. The objective of the DDM is to increase the surface area of the NanoSat in order to decrease the ballistic coefficient of the S/C, which then will reduce the deorbiting time below 25 years (note that, under the assumptions presented in the Mission Analysis section, for a SSO LEO orbit, the DDM will likely only be useful up to 700 km altitude). The idea is to provide a scissor-like deployment of the MLI from its stowed position into the deployed status, like a sail. Special care will be taken to wrap/fold the MLI in a way that avoids any hang-up risk during deployment (BB testing is proposed).

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The information from Mission Analysis is that an increase in size to 50x50 cm will be sufficient. 2 DDM’s shall be used per S/C. The main characteristics can therefore be summarized as follows: • 4 Hinges per DDM (same type as for SDM), HDRM is not part of the DDM • CFRP frame, 1 static and 2 deployable arms • Flexures, “saloon-door”-opening with oscillating over-swings • Hinge flexures for actuation and latch as well • MLI provides surface increase, supplier is www.ruag.com.

Deorbit Deployment Mechanism (DDM) concept for NanoSat

MLI Kapton sail (1x)

CFRP deployable arms (2x)

CFRP lower frame (1x)

Maeva-Hinge (4x)

Figure 8-4: DDM design concept 8.3.2.2 Key Characteristics The DDM key characteristics are: • Surface increase to 50x50 cm size to decrease ballistic coefficient • Mass: 287 g (using same hinges as for the SDM) • Dimensions: 0.3x0.1x0.2 m • Temp.: -75 to 105 degC

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• Power: none. 8.3.2.3 Scalability The DDM is scalable for different MLI surface areas. 8.3.2.4 Multi mission suitability There is no restriction for use on NanoSat missions as long as LEO is used. Note: this mechanism will not perform in GTO, and, as above, will likely only be useful to 700 km altitude in SSO (LEO). 8.3.3 Module 3: Hold-Down and Release Mechanism (HDRM) 8.3.3.1 Description and Overview The main task of the HDRM is to stow moving appendages during launch, or to stow the DDM moving parts during launch, and to release them upon command. The selected devices are based on thermal actuation of a shape memory Ti-Ni alloy, which severs a bolt due to the phase change elongation of the device. The smallest device available to date is the US product Frangibolt FC2 series (see Figure 8-5). In addition to the FC2 a dedicated fastener shall be taken, see Figure 8-6. As the product is subject to ITAR restrictions, it has been decided to implement for NanoSat equivalent European devices, however as they are not existing currently in a comparable design, it is suggested to develop them. This development will be based on the TiNi concepts but will of course avoid any legal interference (patents, trade names, ….). It is therefore considered as TRL2 currently (opposite to the US products with TRL8). The main characteristics can therefore be summarized as follows: • SMA technology, resettable (except fastener) • Used for SDM and DDM only • May be replaced by NanoRelease, TBC • European (non-ITAR, to be developed) component is assumed, based on the heritage item Frangibolt (avoiding applicable patent infringements and trade-names, of course).

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Figure 8-5: Frangibolt FC2-16-31SR2 datasheet (RD[6])

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Figure 8-6: Frangibolt Fastener datasheet (from RD[7])

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8.3.3.2 Key characteristics The key characteristics of the HDRM are: • Mass: 40 g per HDRM (actuator+bolt+structure, SDM needs 2, DDM needs 4) • Dimension: 0.025x0.015x0.015 m • Temp: -65 to 80 degC • Power: 25 W for 20-40 sec, see chart. 8.3.3.2.1 Background on Shape Memory Alloys: (from http://www.tiniaerospace.com/sma.html). SHAPE MEMORY ALLOYS (SMAs) refer to a group of materials which have the ability to return to a predetermined shape when heated. The shape memory effect is caused by a temperature dependent crystal structure. When an SMA is below its phase transformation temperature, it possesses a low yield strength crystallography referred to as Martensite (see Stress-Strain figure). While in this state, the material can be deformed into other shapes with relatively little force. The new shape is retained provided the material is kept below its transformation temperature. When heated above this temperature, the material reverts to its parent structure known as Austenite causing it to return to its original shape (see Phase Transformation figure). This phenomenon can be harnessed to provide a unique and powerful actuator. The most widely used shape memory material is an alloy of Nickel and Titanium called Nitinol. This particular alloy has excellent electrical and mechanical properties, long fatigue life, and high corrosion resistance. As an actuator, it is capable of up to 5% strain and 50,000 psi recovery stress, resulting in ~1 Joule/gm of work output. Nitinol is readily available in the form of wire, rod, and bar stock with transformation temperature in the range of -100° to +100° Celsius. More recently applications in Micro-Electro-Mechanical-Systems (MEMS) has led to the development of Nitinol in the form of sputter deposited thin film.

Stress-Strain Characteristics / Phase Transformation

Figure 8-7: TiNi SMA background (from RD[7]) Properties of Nitinol: • Density 6.45 gm/cm3 • Thermal Conductivity 10 W/moK

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• Specific Heat 322 j/kgoK • Latent Heat 24,200 J/kg • Ultimate Tensile Strength 750-960 MPa • Elongation to Failure 15.5% • Yield Strength (Austenite) 560 MPa • Young's Modulus (Austenite) 75 GPa • Yield Strength (Martensite) 100 MPa • Young's Modulus (Martensite) 28 GPa. 8.3.3.3 Scalability This HDRM is scalable to different sizes, option is to increase/decrease number of items used (typically 2 per SDM or 4 per DDM). 8.3.3.4 Multi mission suitability No restriction for use on S/C, if <80degC (phase transition of the SMA). 8.3.4 Module 4: Nano-Terminator De-Orbit Mechanism (NTDM) 8.3.4.1 Description and Overview The main task of the NTDM is to decrease the deorbiting time in LEO below 25 years at the end of the NanoSat lifetime. However, the currently existing nano-terminator example considered here only works for S/C below 1 kg. The selected devices are based on the drag produced by an passive electrodynamic tether, see Figure 8-8 for more information and description. In addition to the NanoTerminator, the Nanosat Release Actuator (see Figure 8-9) will be included into this module. As the product is subject to ITAR restrictions, it has been decided to implement for NanoSat equivalent European devices, however as they are not existing currently in a comparable design, it is suggested to develop them. This development will be based on the NanoTerminator concepts but will of course avoid any legal interference (patents, trade names, ….). It is therefore considered as TRL2 currently (opposite to the US products with TRL8). The main characteristics can therefore be summarized as follows: • 1 per S/C • Spring-actuated deployment of 100m tether • Including Release actuator PCB • For earth-face mounting • Usable up to 1000km altitude to stay below 25 years deorbit time • Power and activation signal will be provided by the S/C • European (non-ITAR, to be developed) component is assumed, based on the US heritage NanoTerminator and NanoRelease (avoiding applicable patents).

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Figure 8-8: NanoTerminator datasheet (from RD[8]) (Note for above diagram on deorbit time: CubeSat has mass of 1kg and 10 cm cube dimensions).

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Figure 8-9: Nanosat Release Mechanism datasheet (from RD[9])

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8.3.4.2 Key characteristics The key characteristics are: • Mass: 61 g (NanoTerminator + NanoRelease PCB but without Battery) • Dimensions: 0.06x0.04x0.04 m • Temp.: -40 to 85 degC • Power: TBD, but small, to be provided by the S/C. 8.3.4.2.1 Background on Electrodynamic Tether Drag: (from http://www.tethers.com/EDTethers.html) An electrodynamic tether is essentially a long conducting wire extended from a spacecraft. The gravity gradient field (also known as the "tidal force") will tend to orient the tether in a vertical position. If the tether is orbiting around the Earth, it will be crossing the Earth's magnetic field lines at orbital velocity (7-8 km/s). The motion of the conductor across the magnetic field induces a voltage along the length of the tether. This voltage can be up to several hundred volts per kilometre. In an "electrodynamic tether drag" system, such as the Terminator Tether, the tether can be used to reduce the orbit of the spacecraft to which it is attached. If the system has a means for collecting electrons from the ionospheric plasma at one end of the tether and expelling them back into the plasma at the other end of the tether, the voltage can drive a current along the tether. This current will, in turn, interact with the Earth's magnetic field to cause a Lorentz JXB force which will oppose the motion of the tether and whatever it is attached to. This "electrodynamic drag force" will decrease the orbit of the tether and its host spacecraft. Essentially, the tether converts the orbital energy of the host spacecraft into electrical power, which is dissipated as ohmic heating in the tether.

Figure 8-10: Illustration of electrodynamic tether (from RD[9])

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Figure 8-11: Illustration of NanoTerminator (from RD[9]) 8.3.4.3 Scalability One size fits all NanoSats. 8.3.4.4 Multi mission suitability Suitable for 3-axis stabilized and LEO only.

8.4 Scenario Study Case 1 (LEO Without Propulsion) 8.4.1 Design Drivers Provide a lightweight and low-cost mechanism solution for NanoSat. 8.4.2 Module Selection (Justification) Only Module 4 (NTDM) has been selected for the LEO configuration, as the only requirement is to aid with the deorbiting. 8.4.3 Mass Budget

Element 1 NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity incl. margin Click on button above to insert new unit excl. margin

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8.5 Scenario Study Case 2 (LEO With Propulsion) Note, in terms of mechanisms, this configuration is identical to Study Case 1 (LEO without propulsion). 8.5.1 Design Drivers Provide a lightweight and low-cost mechanism solution for NanoSat. 8.5.2 Module Selection (Justification) As with Study Case 1, only Module 4 (NTDM) has been selected. 8.5.3 Mass Budget

Element 1 NanoSat MASS [kg] Unit Unit Name Qu an tity Mass per Maturity Level Margin Total Mass quantity incl. margin Click on button above to insert new unit ex cl. margin

1 S/A Deployment Mechanism SDM (w/o HDRM and S/A) 0 0.086 To be modified 0 0.000 Deorbit Deployment Mechanism (DDM) (w/o HDRM, with sail, 2 not needed if unit-4 is taken) 0 0.287 To be modified 0 0.000 Hold-Dow n and Deployment Mechanism (HDRM, optional, only 3 2 needed if unit 4 is taken) 0 0.040 Fully developed 0 0.000 Nano-Terminator Deorbit Module (NTDM) ( incl nanosat release, 4 would replace 2 DDM and 2 HDRM) 1 0.061 Fully developed 0 0.061 - Click on button below to insert new unit 0.0 To be developed 0 0.0 SUBSYSTEM TO TAL 4 0.061 0.0 0.061 Table 8-3: Mass budget - Case 2 8.5.4 Capabilities Provided to a Payload The NTDM will speed up the de-orbiting at EOL to a duration of less than 25 years. 8.6 Scenario Study Case 3 (GTO) 8.6.1 Design Drivers Provide a lightweight and low-cost mechanism solution for NanoSat. 8.6.2 Module Selection (Justification) The following modules have been selected for NanoSat, to meet the need for deployable solar arrays in this configuration (see the Power section): • SDM • HDRM. As the NTDM (and DDM) are not useful in GTO they are not considered. 8.6.3 Mass Budget

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Element 1 NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity incl. margin Click on button above to insert new unit ex cl. margin 1 S/A Deployment Mechanism SDM (w/o HDRM and S/A) 2 0.086 To be modified 0 0.172 Deorbit Deployment Mechanism (DDM) (w/o HDRM, with sail, 2 not needed if unit-4 is taken) 0 0.287 To be modified 0 0.000 Hold-Down and Deployment Mechanism (HDRM, optional, only 3 2 needed if unit 4 is taken) 4 0.040 Fully developed 0 0.160 Nano-Terminator Deorbit Module ( incl nanosat release, would 4 replace 2 DDM and 2 HDRM) 0 0.061 Fully developed 0 0.000 - Click on button below to insert new unit 0.0 To be developed 0 0.0 SUBSYSTEM TO TAL 4 0.332 0.0 0.332 Table 8-4: Mass budget - GTO 8.6.4 Capabilities Provided to a Payload The SDM with the HDRM will bring two of the S/A panels in an orientation perpendicular to the sun incidence, thus optimising the power output of the solar array.

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9 PROPULSION 9.1 Requirements and Design Drivers The propulsion system modules for the Nano-sized Spacecraft (NanoSat) have been designed to be able to satisfy the following main functions, dependent on the individual mission scenario: • 1 DoF thrust control for orbit manoeuvres or station keeping • 1 DoF thrust control for de-orbit manoeuvres • 3 DoF torque control for attitude control wheel off-loading • 6 DoF thrust and torque control for fine attitude control. The range of possible propulsion related requirements, derived from the mission scenario requirements are: • ∆v budgets: o De-orbit: LEO: 0 GTO: 50 m/s o Orbit injection corrections: LEO: 9.50 m/s GTO: 36 m/s o Target orbit acquisition: LEO: 5 m/s GTO: 0 o Orbit maintenance: LEO: 47 m/s GTO: 0 o AOCS: 6 m/s (typical value) o Orbit changes from GTO: 680 to 1500 m/s (eg. Moon or GEO transfer) 0.016 m/s per km (apogee change) 0.1 m/s / per km (perigee change) 28 m/s per degree (inclination change). • Thrust levels: o Minimum impulse bit: 10 µN o Thrust range of interest: 0.5 mN to 500 mN. • Mission lifetime: 2 years. Other constraints imposed on the propulsion system design were: • Dry mass budget: 1 kg with target 400 to 800 g • Propellant mass budget: 3 kg maximum • Tank size constraint: Ø 200 mm maximum • Power budgets: 10 W for de-orbit, injection and orbit changes; 2 to 3 W for orbit maintenance and AOCS. The following other general requirements were relevant to the NanoSat study:

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• All sub-systems are to be modular and have common interfaces to allow flexibility to meet different mission requirements and to ease integration • Sub-systems shall be scalable, so that use of the same technology shall be possible for a wide range of missions • The proposed sub-systems shall be directly applicable to multiple mission applications • The development status of sub-systems shall be such that TRL 8 is achievable within a 10 year period • Sub-systems shall sit on a digital data bus so all elements must have their own control electronics • Sub-system recurring costs shall be very low. 9.2 Sub-System Description 9.2.1 General The above set of requirements represent a wide range of propulsion needs, and is clearly not achievable with a single propulsion system. The overall propulsion sub-system design was therefore approached by first trying to separate the requirements into groups, aimed at specific propulsion needs (such as orbit maintenance or AOCS wheel off-loading), for which different propulsion modules could be defined. The following generic propulsion needs were identified and classified in terms of the number of DoF, the ∆v and the thrust level: • 1 DoF, high ∆v, mN thrust for LEO de-orbit or GTO orbit changes • 1 DoF, moderate ∆v, mN thrust for orbit injection, acquisition and maintenance • 3 DoF, low or moderate ∆v, µN or mN thrust for AOCS wheel off-loading • 6 DoF, moderate ∆v, µN thrust for fine attitude control. 9.2.2 Technology Trade-off Initially, all possible types of propulsion technologies were considered, and a detailed literature search was performed to identify candidate technologies. This literature search identified the following (non-exhaustive) list of propulsion technologies: • Propellant storage/generation technologies: o Conventional monolithic metal liquid propellant tanks o Novel liquid propellant tank materials (such as plastics) o Conventional composite over-wrapped pressure vessels o Nitrogen gas generators. • Chemical propulsion technologies: o Conventional small mono-propellant (1 N class) o MEMS mono-propellant thrusters o Conventional solenoid valve cold gas thrusters o MEMS cold gas thrusters o Conventional solid fuel rocket motor o MEMS based digital solid fuel thrusters.

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• Electric propulsion technologies: o Conventional resistojets o MEMS based micro-resistojets o Pulsed plasma thrusters o FEEP thrusters o Colloid thrusters o Micro-ion engines. A simple “traffic light” trade-off was performed to identify those technologies which might be suitable to the NanoSat application and those which are clearly unsuitable. The results are presented in the following tables:

Technology System System System System TRL Performance mass / power complex / (dry mass volume requirement cost fraction) Conventional monolithic metal propellant tanks Novel material propellant tanks Conventional composite overwrapped pressure vessels Gas generators

Table 9-1: Propellant storage / generation technology trade-off

Technology System System System System TRL Performance mass / power complex / (Isp) volume requirement cost Conventional small monopropellant thrusters MEMS mono-propellant thrusters Conventional solenoid valve cold gas thrusters MEMS cold gas thrusters Conventional solid fuel rocket motor; MEMS based digital solid fuel thrusters Table 9-2: Chemical propulsion technology trade-off

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Technology System System System System TRL Performance mass / power complex / (Isp) volume requirement cost Conventional resistojets MEMS based micro-resistojets Pulsed plasma thrusters FEEP thrusters Colloid thrusters Micro-ion engines Table 9-3: Electric propulsion technology trade-off The results of the trade off can be summarised as follows: • Conventional monolithic metal liquid propellant storage vessels should be acceptable, but mass benefits might be achievable from considering alternative materials, such as plastics, at the expense of additional development activity • Conventional composite overwrapped pressure vessels are not favoured because of the high tank mass and lack of availability of suitably sized vessels • Gas generators may provide a lower volume propellant storage capability, but the dry mass fraction is extremely poor and the firing system is relatively complex, requiring dedicated electronics • All of the identified chemical propulsion technologies are acceptable, but the conventional solenoid valve cold gas thrusters provide the simplest solution as long as performance demands are moderate • Of the electric propulsion technologies, only pulse plasma thrusters are considered acceptable. All other technologies require too high a power demand and are generally too complex for the NanoSat application. 9.2.3 Generic Sub-System Design Based on the results of the above simple trade-off, a range of propulsion modules were defined, as described in section 9.3. The approach taken was to define separate propellant storage or generation modules, which could be combined with different thruster modules to meet the different mission scenario requirements. In order to keep the propellant tank module simple, only the tank itself, and the necessary fill and drain/vent valves were included in this module (ie. components requiring electronics were excluded). Since the ∆v, and therefore propellant load, required for each mission scenario will be different, it was not possible to size the tank for a specific propellant load. The approach taken was to define a baseline tank size, which might conceivably meet most of the mission needs, and to provide a tank mass vs propellant mass scaling for use in system level budgeting. A gas generator module was defined based on existing technology. In this case the complexity is higher than that of the propellant tank module since a pressure sensor and dedicated electronics are required. Once again the generator could not be sized for a specific propellant load, so a

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 101 of 267 baseline design was defined along with a scaling for dry mass vs propellant mass for system budgeting purposes. Thruster modules based on conventional mono-propellant, cold gas, solid rockets and MEMS based cold gas technologies were defined. In each case the necessary valve, filters and electronics were included in the modules. For the conventional cold gas thruster technology two separate modules were defined with different numbers of thrusters. This was in order to allow multiple thruster modules to be combined with a single propellant tank module to meet different mission needs. No electric propulsion module was defined during the study, but it was recognised that pulsed plasma thrusters could provide significant benefit to this class of mission. This technology should be considered in any future study. Example propulsion systems based on the above described modular approach are presented in Figure 9-1 and Figure 9-2.

Figure 9-1: Proposed Modularisation of Monopropellant Propulsion System

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Cold Gas Generator Module

4 MEMS Thruster Module CG

Figure 9-2: Proposed Modularisation of Cold Gas Generator Propulsion System 9.2.4 Sub-System Interfaces In order to comply with the general requirement for modularity, all the propulsion modules should have the same interfaces. The following standard interfaces were used for the propulsion modules: • Single power bus interface, based on the defined spacecraft power bus • Single communication bus interface, based on the defined spacecraft communications bus • Simple propellant pipework interface (between tank modules and thruster modules) consisting of a single 1/8” stainless steel tube • Simple mechanical bracket for thruster module attachment to the spacecraft. It was not possible to standardize tank or gas generator module mechanical interfaces at this stage. 9.3 Module Descriptions The following list of modules was considered: • Module A: Propellant tank module • Module B: Gas generator module • Module C: Single solid propellant thruster module • Module D: Single mono-propellant thruster module • Module E: Single cold gas propellant thruster module • Module F: Triple cold gas propellant thruster module • Module G: MEMS cold gas thruster pod.

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9.3.1 Module A (Propellant Tank) 9.3.1.1 Description and Overview The propellant tank module is a simple module, consisting of a liquid propellant tank along with the associated mechanical fill and drain valves and a propellant filter. No electrical components are included in the module in order to keep complexity to a minimum. A simple schematic for the module is presented in Figure 9-3.

Figure 9-3: Propellant Tank Module (A) The baseline module consists of the following components: • Tank: ARDÉ D4187 diaphragm tank • Fill and drain valves: EADS Astrium or AMPAC-ISP • Filter: Sofrance or Vacco. 9.3.1.2 Key characteristics 9.3.1.2.1 Tank The baseline tank is from ARDE, number D4187. Key characteristics are listed in Table 9-4 and the tank drawing is presented in Figure 9-4.

Material CRES 310 Expulsion Device Diaphragm Expulsion Efficiency 92 % Diameter [m] 0.207 Volume [l] 3.85 Mass [kg] 1.24 MOP [Bar] 84.6 Table 9-4: ARDÉ D4187 Tank Characteristics

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Figure 9-4: ARDÉ D4187 Tank ICD, RD[11] 9.3.1.2.2 Fill and Drain Valves The baseline fill and drain valves are from EADS Astrium. Key characteristics are listed in Table 9-5 and a photograph of the unit is presented in Figure 9-5. Material CRES 310 Operating Media Hydrazine, NTO, MON Expulsion Efficiency 92 % Tube diameter [mm] 6.35 Length [mm] 112.5 Mass [kg] 0.07 External Leakage [sccs GHe] < 1x10-6

Internal Leakage [sccs GHe] < 2.8x10-4 Table 9-5: Astrium Fill Drain Valve Characteristics, RD[12]

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Figure 9-5: EADS Astrium Fill and Drain Valve 9.3.1.2.3 Filter The baseline filter is from Sofrance, model number RA03473A. Key characteristics and drawings are presented in Figure 9-6.

Figure 9-6: Sofrance RA03473A Filter Characteristics

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9.3.1.3 Module Mass Breakdown The mass breakdown of module A is shown in Table 9-6.

Propellant Mass (kg) 0 102030405060 14

12 [1] Empirical Data

10 [2] Line of Best Fit: Empirical Data

6 Tank Asmbly Mass = 0.1193 (Tank Volume) + 0.55

Tank AssemblyTank Mass (kg) 4 Tank Asmbly Mass = 0.1674 (Prop Mass) + 0.5583 2

0 0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 80.0 90.0 Tank Volume (l)

Figure 9-7: Propellant Tank Sizing Based on Empirical Data 9.3.1.5 Multi mission suitability As discussed above, this module is suitable for all missions requiring storage of liquid propellant.

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9.3.1.6 Development Needs This baseline module is constructed from fully developed components. No significant development is expected. For use with the Butane cold gas thruster modules a diaphragm PMD may not be the optimal solution, since Butane is self pressurising and does not need to be maintained in the liquid form at the tank outlet. Development of PMDs based on a ‘metal foam’ approach should be considered. Additionally, development of lighter tanks using alternative materials (such as plastic) and simplified mounting arrangements should be considered. 9.3.2 Module B (Cold Gas Generator) 9.3.2.1 Description and Overview The cold gas generator module provides Nitrogen propellant for cold gas thruster modules. The module consists of a cold gas generator unit, a pressure transducer and an electronics board (See Figure 9-8). The generator produces the gas through the decomposition of solid material, and hence the propellant storage efficiency may be improved compared to storage of gas in conventional pressurized tanks. Since the cold gas generator provides multiple re-pressurisations of the feed system a pressure transducer is included in the module to monitor pressure levels and notify the unit when to trigger the next gas generator.

Figure 9-8: Cold Gas Generator Module (B) The baseline module consists of the following components: • Cold Gas Generator: TNO • Pressure Transducer: Presens, Kulite or Druk • Electronics: To be developed. 9.3.2.2 Key characteristics 9.3.2.2.1 Cold Gas Generator The baseline cold gas generator is from TNO and has been developed under ESA’s SME-LET initiative. Key characteristics are listed in Table 9-7 and a photograph of the development unit is shown in Figure 9-9.

Size 120 x 70 x 20 mm Unit Mass 130 g Number of charges 12

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Propellant mass per 0.4 g charge Charge pressure 0.1 – 15 MPa Propellant type Nitrogen Table 9-7: TNO Cold Gas Generator Characteristics, RD[13]

Figure 9-9: TNO cold gas generators developed for ESA, RD[13] 9.3.2.2.2 Pressure Transducer The baseline pressure transducer is the 01560 unit manufactured by Presens. Key characteristics are presented in Table 9-8, and a picture and drawing in Figure 9-10.

Calibrated Pressure Range 250 bar Proof Pressure 375 bar Burst Pressure 700 bar Calibrated Temperature Range 0 to 50 °C Acceptance Temperature Range -5 to 55 °C Qualification Temperature -10 to 60 °C Total Pressure Accuracy including non-linearity, hysteresis and < ± 0.2 % FS temperature effects over calibrated temperature range Temperature Measurement Method Pt-100 element, Class B Temperature Measurement Accuracy ± 1 Pressure Port Interface 1/8 inch weldable tube and threaded connection 5/16 × 24 UNF Pressure Port Interface Material AISI316L Housing Material Inconel 625 Supply Voltage 5 to 12 V dc

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Output Signal Analog, - 1 bar 0.5 V ± 0.2 % - 250 bar 4.5 V ± 0.2 % Mass incl. flying leads ~ 0.075 kg Current Consumption at 5 V – 12 V < 5 mA Electrical Interface Flying Leads, 1 m Insulation Resistance between any pin and chassis at 100 V dc. > 100 MΩ Table 9-8: Presens 01560 Pressure Transducer Characteristics, RD[15]

Figure 9-10: Presens 01560 Pressure Transducer, RD[15] 9.3.2.2.3 Electronics An electronics unit will be required to initiate each of the charges in the gas generator, and to provide excitation and signal acquisition for the pressure transducer. The electronics unit will receive power from the 8V spacecraft power bus, and will be connected to the spacecraft communications bus. It is anticipated that the power and communications interfaces will be based on standard developments, to be used for many of the sub-systems. Additional development will be required for the cold gas charge drivers and the pressure transducer driver. 9.3.2.3 Module Mass Breakdown The mass breakdown for module B is shown in Table 9-9.

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 6 Cold Gas Generator Module (B) Tank Assembly 1 0.130 To be developed 0 0.0 7 Cold Gas Generator Module (B) Pressure Transducer 1 0.075 To be developed 0 0.0 8 Cold Gas Generator Module (B) Electronics 1 0.200 To be developed 0 0.0 SUBSYSTEM TOTAL 3 0.405 0.0 0.405 Table 9-9: Module B Mass Breakdown 9.3.2.4 Scalability The existing TNO cold gas generator modules have been developed for a range of applications covering satellite masses in the range 1 kg to 100 kg. As such they already offer a range of propellant charges providing between 0.1 and 40 grams of propellant per charge RD[13]. Table 9-10 shows the range of units that are under development by TNO, in co-operation with

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Bradford Engineering B.V, NL, specifically aimed at micropropulsion applications. The technology is therefore extremely scalable, and could provide a wide range of propellant needs for the Nanosat applications.

Cubesat Nanosat Microsat Cool gas Hot gas Typical satellites 1 kg mass 2 to 10 kg 10 to 100 kg 10 to 100 kg Size 80 x 15 x 40 mm 120 x 70 x 20 mm 50 mm dia, 22 mm dia, 140 mm long 30 mm long Total impulse 0.6 Ns 3.6 Ns 24 Ns - Total mass 120 g 130 g 250 g 10 g (complete system) (gas storage only) (gas storage) (thruster) Plug and play’ yes yes yes no Development Qualifications Demonstration Space qualified Demonstration status expected in 2009 model model Table 9-10: Specifications of Cold Gas Micropropulsion Systems, RD[13] 9.3.2.5 Multi mission suitability As discussed above, this module could be suitable for a range of missions requiring storage of cold gas propellant. The technology can easily be adapted to provide a range of propellant loads depending on mission need. However, the maximum propellant load is limited by the relatively high dry mass of these systems, so that they are only suitable for relatively modest ∆v requirements. 9.3.2.6 Development needs As discussed above, cold gas generators are already under development for a range of spacecraft classes. For any particular Nanosat application it may be necessary to produce a tailored design to meet mission requirements, but this could be based on combining the appropriate number of existing propellant charge units. The electronics unit will be a new development, although power and communication bus interfaces may be based on standard developments for the Nanosat application. 9.3.3 Module C (Single Solid Propellant Thruster) 9.3.3.1 Description and Overview This module is intended to provide a simple propulsion solution for 1 DoF, high ∆v manoeuvres, such as de-orbiting, where accurate thrust control and multiple thruster burns are not required. The module design is based on scaling of existing solid thruster technologies and consists of a fuel casing, nozzle and ignition system. The use of solid fuels has extensive heritage in missile

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 111 of 267 systems and launchers. However, there are no small solid propellant thrusters available which are directly suitable for the NanoSat application. In the US there is an ongoing development by Thiokol/NASA (see Figure 9-11) RD[16], which has been used as the basis for the definition of this module.

Figure 9-11: Prototype Solid Rocket Motor (Thiokol/NASA), RD[16] 9.3.3.2 Key characteristics Based on data from the Thiokol/NASA development, and a NanoSat ∆v requirement of 50 m/s for de-orbit, the key characteristics of the module are presented in Table 9-11.

Size Ø 50 mm x 340 mm Module Wet Mass 675 g Average Thrust 220 N Specific Impulse 260 s Propellant Mass 395 g Total Impulse 2600 Ns Table 9-11: Characteristics of the NanoSat solid propellant thruster 9.3.3.3 Module Mass Breakdown The mass breakdown of module C is shown in Table 9-12.

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 9 Solid Rocket Module (C) Motor 1 0.475 To be developed 0 0.0 10 Solid Rocket Module (C) Nozzle 1 0.100 To be developed 0 0.0 11 Solid Rocket Module (C) Ignition/Electronics 1 0.100 To be developed 0 0.0 SUBSYSTEM TOTAL 3 0.675 0.0 0.675 Table 9-12: Module C Mass Breakdown 9.3.3.4 Scalability A solid propellant thruster of this type is scalable over a wide range of thrust levels and total impulses. Thruster designs could be made to cover a wide range of mission requirements. However, each new mission requirements would need a specific nozzle and fuel casing design.

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9.3.3.5 Multi mission suitability Although it provides a relatively low mass solution to high ∆v manoeuvres, the solid propellant thruster module has limited multi mission suitability. The “one-shot” nature of the device means that it is only suitable for single burn manoeuvres where the absolute magnitude of the required ∆v is not critical. 9.3.3.6 Development Plan A complete new development will be required for this module. 9.3.4 Module D (Single Mono-Propellant Thruster Module) 9.3.4.1 Description and Overview The Single Mono-Propellant Thruster Module is aimed at providing 1 DoF, moderate ∆v manoeuvres, such as orbit insertion or orbit changes. It consists of a single sub-Newton monopropellant thruster, a latch valve, an electronics board, and associated pipework and wiring. A simple schematic for the module is presented in Figure 9-12.

Figure 9-12: Single Thruster Monoprop Module (D) The baseline module design consists of the following components: • Thruster: 0.5 N, EADS Astrium • Latch valve: MOOG 52-266 • Electronics board: To be developed. 9.3.4.2 Key characteristics 9.3.4.2.1 Thruster The baseline thruster is a 0.5 N Hydrazine thruster manufactured by EADS Astrium. Key characteristics are listed in Table 9-13, and the thruster is shown in Figure 9-13.

Propellant Hydrazine Thrust (in vacuum) 0.5 N Thrust Power 0.55 kW 0.75 hp Isp (in vacuum) 227.3 s Chamber Pressure 22 bar Overall Length 113 mm

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Nozzle Diameter 4.8 mm Mass 0.195 kg

Table 9-13: Astrium 0.5 N Hydrazine Thruster Model CHT 0.5 Characteristics, RD[17]

Figure 9-13: Astrium 0.5 N Hydrazine Thruster Model CHT 0.5 9.3.4.2.2 Valve The baseline latch valve is manufactured by MOOG. Key characteristics are listed in Table 9-14 and a picture and drawing of the valve are shown in Figure 9-14.

Pressures: - Operating 0 to 24 bar - Proof 72.12 bar Operating Voltage Range 21 to 32 Vdc Response Time < 50 msec with 30 Ω coils at 24 bar Power Consumption ≤ 16 watts at 21.5 Vdc at 21°C Leakage: - Internal ≤ 1×10-3 scc/sec GHe at 2.7 to 24 bar - External ≤ 1×10-6 scc/sec GHe at 24 bar Cycle Life ≥ 20,000 Weight 0.65 kg Reverse Cracking Pressure 1.72 to 13.79 bar Inlet Filtration 45 micron absolute with integral filter Thermal Capability (operating) -3.9°C to + 55°C Materials of Construction Titanium, Teflon ® Table 9-14: MOOG Latch Valve 52-266, RD[18]

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Figure 9-14: Photo and Installation Drawing for Latch Valve. All dims in Inches (cm). RD[18] 9.3.4.2.3 Electronics An electronics unit will be required to drive the latch valve, and thruster valves and heater. The electronics unit will receive power from the 8V spacecraft power bus, and will be connected to the spacecraft communications bus. It is anticipated that the power and communications interfaces will be based on standard developments, to be used for many of the sub-systems. Additional development will be required for the valve and thruster drivers. 9.3.4.3 Mass Breakdown In addition to the thruster and latch valve further piping and wiring are used within the module. The mass breakdown of module D is shown in Table 9-15.

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 12 1 Thruster Monoprop Module (D) Thruster 1 0.195 To be developed 0 0.195 13 1 Thruster Monoprop Module (D) Latch Valve 1 0.650 To be developed 0 0.650 14 1 Thruster Monoprop Module (D) Electronics 1 0.100 To be developed 0 0.100 15 1 Thruster Monoprop Module (D) Piping 1 0.050 To be developed 0 0.050 16 1 Thruster Monoprop Module (D) Wiring 1 0.050 To be developed 0 0.050 SUBSYSTEM TOTAL 5 1.045 0.0 1.045 Table 9-15: Module D Mass Breakdown 9.3.4.4 Scalability This module is based on existing qualified propulsion components, so the performance, in terms of thrust and specific impulse is fixed. The total impulse that can be provided depends on the amount of propellant which can be loaded into the tank. Scalability to higher thrust levels would be possible by replacing the selected thruster with a larger unit (1 N and 10 N devices exist), but scalability to lower thrust levels would require the development of a new smaller thruster.

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9.3.4.5 Multi mission suitability This thruster module was designed to provide 1 DoF, moderate ∆v manoeuvres. It is suitable for any mission requiring this type of propulsion manoeuvre. 9.3.4.6 Development needs The baseline module is predominantly constructed from qualified components. No significant development is expected for the thruster and valves, although development of smaller, lower mass European valves and thrusters would be advantageous. As for many of the other propulsion modules the electronics unit will be a new development, although power and communication bus interfaces may be based on standard developments for the Nanosat application. 9.3.5 Module E (Single Cold Gas Thruster) & Module F (Triple Cold Gas Thruster) 9.3.5.1 Description and Overview The Single Cold Gas Thruster Module is aimed at providing 1 DoF, moderate ∆v manoeuvres, such as orbit insertion or orbit changes. It consists of a single milli-Newton class thruster valve, a propellant vaporizer, an electronics board, and associated pipework and wiring. A simple schematic for the module is presented in Figure 9-15. Use of a number of the Triple Cold Gas Thruster Module is aimed at providing 3 DoF, low ∆v manoeuvres for AOCS wheel off-loading. The triple cold gas thruster module will consist of three thrusters, a vaporiser, an electronics board, and associated pipework and wiring. A simple schematic for the module is presented in Figure 9-16. Butane is selected as the propellant due to its storability, enabling a low pressure liquid propellant tank to be used rather than a high pressure gas tank. This choice also reduces the complexity of the system since a gas pressure regulator will not be required, although a propellant vaporizer may be needed to ensure no liquid propellant reaches the thruster valve.

Figure 9-15: Single Cold Gas Thruster Module Figure 9-16: Triple Cold Gas Thruster Module (E) (F) The baseline module consists of the following components: • Thrusters: AMPAC-ISP SV14 or equivalent • Vaporiser: New development, could be based on AMPAC-ISP TT-01 • Electronics board: To be developed.

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9.3.5.2 Key characteristics 9.3.5.2.1 Thruster The baseline thruster is an AMPAC-ISP SV14 unit which has been used on a number of missions including Cryosat. Key characteristics are listed in Table 9-16 and pictures are shown in Figure 9-17.

Operational Temp -35°C to +65°C Operating Pressure 2.5 bar Vacuum Thrust 10 to 40 mN (± 5%) Specific Impulse > 60 s (on Butane) Coil Resistance 105 – 115 Ω at 20°C Power Consumption < 3.5 W (pull-in) < 0.7 W (holding) Response Time < 4.0 ms (opening & closing) Mass < 0.075 kg Maximum O/D 15.85 mm Cycle Life 2,000,000 External Leakage ≤ 1×10-6 scc/sec GHe at 1.5 bar ≤ 2.7×10-4 scc/sec GHe at 1.5 bar Internal Leakage Construction: - Body & Magnetic Stainless Steel parts 304L/Radiometal 4550 - Seal Silicone Rubber Mechanical Interface - Inlet 7/16-20 UNJF-3A - Mounting 37° STD Flare – Male 4 off M3 holes on 23 mm PCD Maximum Length 52 mm Operating Fluid GN2, Xe, Butane Table 9-16: AMPAC-ISP Cold Gas Thruster Model SV14 Characteristics, RD[19]

Figure 9-17: AMPAC-ISP Cold Gas Thruster Model SV14

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9.3.5.2.2 Vaporiser The baseline vaporiser will be a new development, but could be based on the AMPAC-ISP TT- 01 thermothrottle design. Key characteristics are listed in Table 9-17 and drawings are shown in Figure 9-18.

Operating Media GHe, GN2, GXe, Dry air Operational Temp -40°C to +100°C Operating Pressure 2.5 bar Power 0 to 8 W Heater Resistance 4 Ω @ 20°C (Typical) Response Time Flow reduction < 20 s Flow increase < 60 s External Leakage < 1×10-6 scc/sec GHe Flow Range 0.05 mg/s to 20 mg/s (Typical) Cycle Life 7,000 Materials Stainless Steel, Radiometal, Ceramic Mass < 0.020 kg Pipe Connections 1/16” OD Stainless Steel Tube Housing Dimensions 24.5 × 12 × 19 mm Table 9-17: AMPAC-ISP Thermal Throttle TT01 Characteristics, RD[20] R

Figure 9-18: AMPAC-ISP TT01 Thermal Throttle 9.3.5.2.3 Electronics An electronics unit will be required to drive the thruster valves and vaporizer heater. The electronics unit will receive power from the 8V spacecraft power bus, and will be connected to the spacecraft communications bus. It is anticipated that the power and communications interfaces will be based on standard developments, to be used for many of the sub-systems. Additional development will be required for the thruster valve and thermothrottle drivers.

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9.3.5.3 Mass Breakdown In addition to the thruster and vaporiser, further piping and wiring are used within the modules. The mass breakdown of module E is shown in Table 9-18, and the mass breakdown of module F is shown in Table 9-19.

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 17 1 Thruster Butane Module (E) Thruster 1 0.075 To be developed 0 0.075 18 1 Thruster Butane Module (E) Vaporiser 1 0.050 To be developed 0 0.050 19 1 Thruster Butane Module (E) Electronics 1 0.100 To be developed 0 0.100 20 1 Thruster Butane Module (E) Piping 1 0.050 To be developed 0 0.050 21 1 Thruster Butane Module (E) Wiring 1 0.050 To be developed 0 0.050 SUBSYSTEM TOTAL 5 0.325 0.0 0.325 Table 9-18: Module E Mass Breakdown

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 22 3 Thruster Butane Module (F) Thruster 1 0.225 To be developed 0 0.225 23 3 Thruster Butane Module (F) Vaporiser 1 0.050 To be developed 0 0.050 24 3 Thruster Butane Module (F) Electronics 1 0.150 To be developed 0 0.150 25 3 Thruster Butane Module (F) Piping 1 0.150 To be developed 0 0.150 26 3 Thruster Butane Module (F) Wiring 1 0.150 To be developed 0 0.150 SUBSYSTEM TOTAL 5 0.725 0.0 0.725 Table 9-19: Module F Mass Breakdown 9.3.5.4 Scalability These modules are based on existing qualified cold gas thrusters, so basic dimensions and the performance, in terms of thrust and specific impulse is fixed. However, scalability to lower or higher thrust levels is possible by redesigning the nozzle section of the thruster valve. The total impulse that can be provided depends on the amount of propellant which can be loaded into the tank. 9.3.5.5 Multi mission suitability The intention of these thruster modules was to provide moderate amounts of ∆v’s in 1 or 3 DoF. The use of a combination of these modules should allow a range of mission requirements to be met, ranging from 1 DoF orbit maintenance manoeuvres through to full 6 DoF attitude control. 9.3.5.6 Development needs Although the baseline module is based on a fully developed thruster valve, this unit has been designed to be powered from a 28 V bus. Some redesign and development of the solenoid coil will be required to make it compatible with the NanoSat 8 V bus. As stated above under scalability, if different thrust levels are required some development will be needed to redesign the nozzle. A qualified Butane propellant vaporizer does not exist, so this unit will need developing. This development could be based on existing thermothrottle technology, such as the AMPAC-ISP

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TT01. These are relatively simple heater devices aimed at changing the flow rate of gases such as Xenon by changing the temperature of the gas as it flows through a restrictive flow path. The same approach could be used to ensure only gaseous Butane reaches the thruster valve. As for previous modules the electronics unit will be a new development, although power and communication bus interfaces may be based on standard developments for the Nanosat application. 9.3.6 Module G (MEMS Cold Gas Thruster Pod) 9.3.6.1 Description and Overview This module is aimed at providing the NanoSat mission with a low, variable thrust capability to be used when accurate attitude control is needed. A combination of 4 of these thruster pods would be able to offer full 6 DoF thrust and torque control for fine attitude control. The thruster pods have been developed by NanoSpace for small satellite and formation flying applications, and will be flight demonstrated on the PRISMA spacecraft. The pod consists of a conventional metal housing in which a stack of silicon wafers is mounted. These wafers provide all the functionality of four cold gas thrusters including filters, isolation valves, heaters and proportional thruster valves. In addition to the thruster pods, the module will need an electronics board capable of interfacing with the spacecraft and providing drivers for the four thrusters. A schematic of the module is shown in Figure 9-19. PT Filter Comms Bus Electronics Board Power Bus MEMS Thruster Pod

Figure 9-19: Four Thruster MEMS Nitrogen CG Module (G) The baseline module consists of the following components: • MEMS thruster pod: Nanospace • Filter: Integrated in thruster pod • Electronics: To be developed. 9.3.6.2 Key characteristics 9.3.6.2.1 MEMS Thruster Pod The baseline thruster pod is developed by NanoSpace and its key characteristics are listed in Table 9-20 and a picture is shown in Figure 9-20.

Propellant Nitrogen Specific Impulse 70 s Thrust range 1 – 1000 µN

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Power Consumption < 1W (per thruster) Dimensions of Housing: - Diameter 43.5 mm - Height 51.0 mm Table 9-20: Thruster Pod Characteristics, RD[22]

Figure 9-20: MEMS Thruster Pod by Nanospace 9.3.6.3 Mass breakdown In addition to the MEMS thruster pod further piping and wiring are used within the module. The mass breakdown of module G is shown in Table 9-21.

Element 1 NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 27 4 Thruster MEMS Nitrogen CG Module (G) Thuster Pod 1 0.120 To be developed 0 0.1 29 4 Thruster MEMS Nitrogen CG Module (G) Electronics 1 0.200 To be developed 0 0.2 30 4 Thruster MEMS Nitrogen CG Module (G) Piping & Wiring 1 0.051 To be developed 0 0.1 SUBSYSTEM TOTAL 30 0.371 0.0 0.4 Table 9-21: Module G Mass Breakdown 9.3.6.4 Scalability The four thruster MEMS module has been designed to provide fine thrust control in four directions. The thrust level of these devices is inherently scalable between 1 to 1000 micro Newtons. Scaling to higher thrust levels would be possible, but would require development of a new MEMS wafer stack. 9.3.6.5 Multi mission suitability This module is only applicable to missions requiring accurate, variable thrust and torque control. 9.3.6.6 Development needs The thruster pods are at a high TRL, having been developed for a flight demonstration on the PRISMA spacecraft. However, a full qualification has not been performed and, as for other propulsion modules, a full development of the electronics board will be required.

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9.4 Scenario Study Case 1 (LEO SSO, 3-Axis Stabilised, Without Propulsion) On-board propulsion is not necessary in this scenario. 9.5 Scenario Study Case 2 (LEO SSO, 3-Axis Stabilised, With Propulsion) 9.5.1 Design Drivers In this scenario on-board propulsion is required for 1 DoF orbit acquisition and maintenance. The design drivers for this scenario, taken from the requirements presented in section 9.1, are: • Orbit injection corrections: 9.5 m/s • Target orbit acquisition: 5 m/s • Orbit maintenance: 47 m/s • Thrust range of interest: 50 mN to 500 mN • Mass budget (wet): 4 kg • Tank size constraint: Ø 200 mm maximum. 9.5.2 Module Selection (Justification) The key design driver for this scenario is the relatively large ∆v required for the orbit maintenance manoeuvres. In addition, a reasonable thrust authority is required in order to maintain short propulsion manoeuvre times. Since the solid thruster and multiple thruster modules are not applicable to this scenario, the following thruster modules were considered: • Module D: Single mono-propellant thruster • Module E: Single cold gas propellant thruster. The higher performance of the mono-propellant option offers an increase in the available ∆v for a given propellant load. However, the additional mass and complexity of the hardware, and issues around the added complexity of propellant handling make this option less attractive. The butane cold gas thruster option, although having significantly lower specific impulse, can provide the required ∆v within the wet mass and tank size limits. This option was therefore selected. Module A: Propellant tank module, was selected to provide the necessary liquid propellant storage. 9.5.3 Mass/Power Budget The dry mass budget for this scenario case is presented in Table 9-22.

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NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 1 Propellant Tank Module (A) Tank 1 1.247 To be developed 0 1.247 2, 3 Propellant Tank Module (A) Fill/Vent Valve 2 0.070 To be developed 0 0.070 4 Propellant Tank Module (A) Filter 1 0.100 To be developed 0 0.100 17 1 Thruster Butane Module (E) Thruster 1 0.075 To be developed 0 0.075 18 1 Thruster Butane Module (E) Vaporiser 1 0.050 To be developed 0 0.050 19 1 Thruster Butane Module (E) Electronics 1 0.100 To be developed 0 0.100 20 1 Thruster Butane Module (E) Piping 1 0.050 To be developed 0 0.050 21 1 Thruster Butane Module (E) Wiring 1 0.050 To be developed 0 0.050 SUBSYSTEM TOTAL 9 1.742 0.0 1.812 Table 9-22: Mass Budget Scenario Study Case 2 The maximum Butane propellant mass that can be stored in the 200 mm diameter tank is 1.99 kg. This mass of propellant is sufficient to provide a ∆v of 62.5 m/s, assuming a spacecraft wet mass of 20 kg, which provides a 1 m/s margin over the requirement of 61.5 m/s. The power budget for the propulsion system is difficult to define at propulsion sub-system level, as it depends on the thruster operating duty cycle, which must be defined by the mission. Typical powers required by the thruster during on-cycles are as follows: • Pull in (30 ms duration): 3.5 W • Holding: 0.7 W. 9.5.4 Capabilities Provided to a Payload With this combination of modules a ∆v of 62.5 m/s can be provided, assuming a spacecraft wet mass of 20 kg. This gives a 1 m/s margin over the requirement of 61.5 m/s. 9.6 Scenario Study Case 3 (GTO, Inertial Pointing Observatory) 9.6.1 Design Drivers In this scenario on-board propulsion is required for 1 DoF orbit injection correction and de-orbit manoeuvres, and 3 DoF AOCS wheel off-loading manoeuvres. The design drivers for this scenario, taken from the requirements presented in section 9.1, are: • Orbit injection corrections: 36 m/s • De-orbit burn: 50 m/s • AOCS: 6 m/s • Thrust range of interest: 50 mN to 500 mN • Mass budget (wet): 4 kg • Tank size constraint: Ø 200 mm maximum. 9.6.2 Module Selection (Justification) In order to provide both 1 DoF and 3 DoF manoeuvres, and achieve the required amounts of ∆v’s, it is necessary to embark a number of different propulsion modules for this scenario study case. Once again the overall ∆v requirement is relatively large for the size of the spacecraft. In particular, the ∆v for de-orbit is a large fraction of the overall, and would consume a significant amount of propellant in the case that a Butane cold gas system was selected for this function.

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The orbit injection correction and AOCS wheel off-loading ∆v’s are within the capability of the Butane cold gas system. So, considering the additional complexity of implementing a monopropellant system, the Butane cold gas system was selected for these manoeuvres. A combination of two triple thruster modules (Module F) and one single thruster module (Module E) was selected. Module A: Propellant tank module, was selected to provide the necessary liquid propellant storage. Module C: Single solid propellant thruster was selected to perform the de-orbit burn manoeuvre. This is a “one shot” manoeuvre which does not require accurate control, so that the solid thruster solution is ideally suited. 9.6.3 Mass/Power Budget

NanoSat MASS [kg] Unit Unit Name Quantity Mass per Maturity Level Margin Total Mass quantity excl. incl. margin margin 1 Propellant Tank Module (A) Tank 1 1.247 To be developed 0 1.247 2, 3 Propellant Tank Module (A) Fill/Vent Valve 2 0.070 To be developed 0 0.070 4 Propellant Tank Module (A) Filter 1 0.100 To be developed 0 0.100 9 Solid Rocket Module (C) Motor 1 0.475 To be developed 0 0.000 10 Solid Rocket Module (C) Nozzle 1 0.100 To be developed 0 0.000 11 Solid Rocket Module (C) Ignition/Electronics 1 0.100 To be developed 0 0.000 17 1 Thruster Butane Module (E) Thruster 1 0.075 To be developed 0 0.075 18 1 Thruster Butane Module (E) Vaporiser 1 0.050 To be developed 0 0.050 19 1 Thruster Butane Module (E) Electronics 1 0.100 To be developed 0 0.100 20 1 Thruster Butane Module (E) Piping 1 0.050 To be developed 0 0.050 21 1 Thruster Butane Module (E) Wiring 1 0.050 To be developed 0 0.050 22 3 Thruster Butane Module (F) Thruster 2 0.225 To be developed 0 0.0 23 3 Thruster Butane Module (F) Vaporiser 2 0.050 To be developed 0 0.0 24 3 Thruster Butane Module (F) Electronics 2 0.150 To be developed 0 0.0 25 3 Thruster Butane Module (F) Piping 2 0.150 To be developed 0 0.0 26 3 Thruster Butane Module (F) Wiring 2 0.150 To be developed 0 0.0 SUBSYSTEM TOTAL 22 3.142 0.0 3.937 Table 9-23: Mass Budget Scenario Study Case 3 The maximum Butane propellant mass that can be stored in the 200 mm diameter tank is 1.99 kg. This mass of propellant is sufficient to provide a ∆v of 62.5 m/s, assuming a spacecraft wet mass of 20 kg, which provides a 20.5 m/s margin over the requirement of 42 m/s. The propellant mass required to provide 42 m/s, using the same assumption on spacecraft wet mass is 1.36 kg. The power budget for the propulsion system is difficult to define at propulsion sub-system level, as it depends on the thruster operating duty cycle, which must be defined by the mission. Typical powers required by the Butane thrusters during on-cycles are as follows: • Pull in (30 ms duration): 3.5 W • Holding: 0.7 W. A short iginition pulse will also be required for the solid thruster, but this will not be a significant issue for the power budget. 9.6.4 Capabilities Provided to a Payload This combination of propulsion modules can provide up to 62.5 m/s for orbit and attitude control manoeuvres in either 1 DoF (for orbit corrections or maintenance) or 3 DoF (for wheel off- loading). In addition the solid thruster provides a 50 m/s capability for de-orbit burn.

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10 POWER 10.1 Requirements and Design Drivers This feasibility assessment investigates how far it is possible to reduce the mass and size of a highly modular, multipurpose power subsystem, which is capable of providing sufficient onboard energy to a payload in multiple scenarios, when applying disruptive technologies and extreme miniaturization. The power subsystem also needs to have a low recurring cost and be readily configurable for a large range of potential missions and payloads. It is expected that, to achieve this, the subsystem design concept needs to be mostly single string (redundancy to be provided at mission level if needed). Further, it is expected that a high amount of technology development effort will be required, firstly to shrink the mass/volume without affecting performance, and secondly to come to designs that have very low recurring costs. Modularity of design is expected to be essential to ensure the configurability of the power subsystem to different missions. With these strong and difficult constraints a high development cost and long development time is expected. No "off the shelf" equipment exists to satisfy these Study objectives and therefore feasible technology developments and equipment designs are identified. Finally, considering a satellite mass between 10 and 20 kg, the mass allocation for the power subsystem is 30%, i.e. between 3.3 and 6.6 kg. 10.1.1 Power Subsystem Modularization As with other subsystems, power subsystem can be seen as a set of preconfigured modules ready to be combined according to mission and configuration needs. Such modules shall be simple, with standard interfaces, in order to allow a massive production with consequent reduction of recurring costs. In order to save mass, the power system architecture shall be battery regulated, with the battery directly connected to the power bus, without the use of Battery Charge and Discharge Regulators (BCDR) A set of modules comprising 2 solar panels, 1 power converter, and 1 battery shall be capable to provide 30W in sunlight (sun-pointing), with an allowed voltage drop of 1V during discharge of the battery in eclipse. Finally, the power distribution shall be defined as integrated with the discrete command bus with the objective of a dramatic reduction of harness and further electrical integration activities (plug- and-play) approach.

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10.2 Sub-System Description

Figure 10-1: Power Subsystem Layout The power subsystem architecture is depicted in Figure 10-1. It comprises the following modules: • Solar panels (S/P #1,…, S/P#n) • Solar Power Conditioners, based on MPPT super-buck topology • Battery Pack. The Po/CB is the Power and Command Bus. Each solar array, either body mounted or deployed, is connected to 1 MPPT module, in this way it is always possible to extract the maximum available power from the solar array in any temperature and illumination condition, and regardless of the solar array positioning with respect to the satellite body. Up to two solar panel modules can be connected to the same converter, according to the overall satellite configuration needs. The overall available power is collected to a common bus which voltage is generally dictated by the battery level. Nevertheless the nominal bus voltage is 8 V. Such value has been decided taking into account three different considerations: • The Li-Ion cell voltage at full charge is 4.2 V • The requirement of 30 W minimum capability implies a maximum bus current of about 8 A with a power bus at 4 V, therefore a 8 V bus is beneficial for the harness mass (current reduced by one half)

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• The modular power bus has to be suitable for all kinds of missions, including LEO and GEO missions, concerning the latter; it has to be observed the absence of eclipse for several months, during such periods the battery cannot be kept fully charged because of progressive capacity fading. The DCM is the module integrating the OBC and the power bus controller. The discrete commands (ON/OFF/RESET) generated by OBC are routed on the power bus as sequence of bits according to a simple protocol. The DCM has then been divided into two parts: a digital one, and a power conditioning and distribution one. The bus voltage control has to be centralised for all the MPPTs and it is not advisable to have voltage control signals circulating in discrete wire outside the units. For this reason, the MPPT modules must be mechanically integrated with the DCM. Therefore the power part of the DCM will provide the bus voltage control, the battery connection, and the main power bus protection. As it can be seen in Figure 10-1, the power bus is now a common rail to which all the loads are connected and they receive both electrical power and discrete commands. This architecture means that a traditional centralised module for the power bus protections is not feasible. The proposed solution is given in Figure 10-2.

Figure 10-2: Proposed power bus protection Note the main power bus has one single protection in the DCM; such protection is activated in case of failure in the main power line, which in turn is double isolated. Latching Current Limiters are then to be positioned at the input of each unit load. They will protect the main power line from local short circuits. Clearly the current limitation capability of the main bus LCL will be higher than the capabilities of the peripheral limiters; as well as its trip-off time will be longer, so that the local LCLs can intervene in case of local short without triggering the main one (selectivity of protections). 10.2.1 Interfaces Interfaces with other subsystems; the power subsystem has a main interface with the DHS within the DCM; this is implicit in the proposed architecture because of the use of the power rail for the transmission of the discrete commands to the different units. The interface with each single unit is composed of a derivation from the power line followed by the local protection LCL. Such LCL can be either a device developed separately from the load unit and then plugged at the

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 128 of 267 integration, or a device developed by the unit’s manufacturer in compliance with the well defined interface requirements. In the first case the power subcontractor shall be responsible for the power part of the DCM, the power line, and the local protections. Interfaces between the modules: each MPPT module will accept up to two solar panels connected at its input. For the reasons said above, each MPPT module shall be mechanically connected with the other MPPT modules and with the rest of the DCM by a unique flexible PCB (printed circuit board). Therefore mechanical compatibility of MPPT modules with DCM modules must be assured. The power part of the DCM shall contain the battery input and the power rail output. 10.3 Module Descriptions 10.3.1 Module 1: Solar Panel 10.3.1.1 Description and Overview The solar panel module is based on the adoption of thin film solar cells which it is foreseen to have available within 10 years, the estimated efficiency of such cells is 26% with a mass of about 1g. The current solar array designs are based on the adoption of triple junction solar cells with 28% efficiency, and a mass of about 2.6g. The panel will be composed of a number of strings connected in parallel and capable to deliver at least 15W EOL. The number of cells to be placed in series for each string has to be a function of the power bus voltage and the power converter topology (super-buck, in our case). In fact the minimum output voltage of the solar array has to be higher than the bus one. The difference between these two voltages is calculated considering the PWM converter switching frequency and the duty ratio: V D = out Vin For a switching frequency of about 100 kHz, the minimum output voltage cannot be below 9.2V. Taking into consideration the main parameters of such cells:

@28 °C @ 100 °C

ISC [MA] 507 507

IMP [MA] 491 491

VMP [V] 2.2 1.876

VOC [V] 2.45 2.126

Table 10-1: Main parameters of solar cell The current is basically considered as unaffected by the temperature, while the voltage decrease is -4.5 mV/°C. Therefore a single string has to be composed of 5 cells connected in series. The number of string in parallel will be such that the output power cannot be lower than 15W; considering 26% efficiency, with an illumination of 1353 W/m2 in winter solstice, the output

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 129 of 267 power of a single cell is 1.06 W. Hence there must be 4 strings. Such configuration: 5s4p guarantees 16 W output. The reduced dimensions of the panel, due to the low number of cells, will implicitly be beneficial for the stiffness of the panel itself because the natural frequencies of it will be higher than those generated by the launcher, this allows for a reduction of the thickness of the substrate honeycomb; then, concerning the mass estimation a conservative figure of 2 kg/m2 for the honeycomb density is assumed. 10.3.1.2 Key characteristics Each solar cell has the following dimensions 80x40 mm, thickness: 20 µm, cover-glass thickness: 50 µm. Panel dimensions: 340x230x10 mm. Panel Mass: 176 g. Delivered power: 16 W. 10.3.1.3 Scalability Each solar panel module provides one half of the minimum required power, therefore two panels mounted with the same orientation can satisfy the demand, the Super-buck MPPT converter can then manage the power of the two panels together. 10.3.1.4 Multi mission suitability The panel in itself is a “brick” to be used in the formation of larger panels if necessary; such need is defined by the mission profile (whichever orbit and attitude) and the power demand from the loads. 10.3.2 Module 2: Solar Array Converter (MPPT) 10.3.2.1 Description and Overview The power conditioning module is based on super-buck (low ripple) PWM converter topology, such topology is nowadays widely used in European space missions in conjunction with MPPT (Maximum Power Point Tracker) control electronics, in order to exploit the maximum power provided by the photovoltaic panels in all illumination and temperature conditions. At the moment there are on the market power converters capable to condition the required power (30 W) having a mass of 120 g. However one of the biggest issues in the development of miniaturised DC-DC converters is related to the implementation of the inductors; with respect to this, different technologies are being considered for the implementation of planar inductors, and among them the LTCC (Low Temperature Co-fired Ceramics) is quite promising. Such technology is used for the fabrication of the inductor cores. Figure 10-3 shows the dimensions of such core for the inductor of a buck converter capable to deliver 10 A at 1 V output having 5 V at input. The diameter is less than 1 cm for a power capability of 20 W.

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Figure 10-3: LTCC Inductor mounted on test PCB More information about this can be found in RD[23]. However, from these results it can be inferred that within 5 years a buck converter with 30 W power capability, and a MPPT controller based on ASIC technology is feasible. The bibliographic reference does not report any detailed information about mass or converter efficiency; therefore the characteristics considered in the following section remain conservative and based on the available technology available today, provided that such characteristics are already in line with the initial system requirements indicating 30% of satellite mass allocation for the power subsystem. 10.3.2.2 Key Characteristics Each power converter has a mass of about 120 g, as said above, the dimensions of the module has to be compatible with those of the DCM. The delivered power is 32 W (for two SA panels). 10.3.2.3 Scalability Each Super-buck MPPT converter module can manage the power of the two panels together, provided that those panels operate under the same illumination and temperature conditions. 10.3.2.4 Multi mission suitability The converter module in itself is a “brick” to be used in the connection of different solar panels with different illumination and temperature conditions, to a common power bus bar. Such need is defined by the mission profile (whichever orbit and attitude) and the power demand from the loads.

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10.3.3 Module 3: Battery 10.3.3.1 Description and Overview The battery module is base on Li-Ion technology. Currently used cells have an energy density of about 130 Wh/kg. Such energy density can be in the future increased by the use of nanotubes technology for the implementation of electrodes. Such innovation can lead to an increase up to 30%. The basic module is composed by 6 cells: 2s3p, capable to provide 20 W for 35 minutes eclipse for a bus voltage decrease of about 1 volt. The dimension and mass of the module are calculated having the Sony 18650HC cell as reference, with a further mass reduction of the cell assembly by 30%. 10.3.3.2 Key Characteristics Configuration: 2s3p, mass: 270 g caging inclusive, dimensions: 4.6 x 6.4 x 6.5 cm. The allowed voltage decrease in discharge (eclipse) is 1 V. 10.3.3.3 Scalability Several modules can be connected in parallel, according to the mission needs; the main requirement to be respected is always 1V drop in the power bus at the end of eclipse. 10.3.3.4 Multi mission suitability See section above. 10.4 Scenario Study Case 1: (LEO w/o Propulsion) 10.4.1 Design Drivers The LEO mission is characterised by a power demand of about 15 W averaged in both sunlight and eclipse. The solar panels are all body mounted as illustrated in the Configuration section, and it is clear that a nadir pointing attitude requires more than two panels (16 W each) because of the non constant illumination of them during sunlight. Table 10-2 below reports the applicable power budget.

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Table 10-2: LEO Power Budget 10.4.2 Module Selection (Justification) The modules for this mission are: • 5 solar panels body mounted • 5 MPPT solar array converters • 1 Battery. The following simulation results justify the need of the chosen configuration of the solar panels in order to provide the correct energy to the platform and the battery recharge. Two mission scenarios have been simulated for a sun-synchronous orbit. The first one is a noon-midnight orbit, concerning the maximum duration of the eclipse. The second one is 9:00 – 18:00 Hr orbit, just for verification of the available energy budget.

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Noon-midnight orbit

Figure 10-4: LEO Mission Energy Budget (12:00 – 24:00 hrs Orbit) Only four panels are considered, as the top panel is assumed not to provide any tangible output in this configuration. A further simulation over several orbits shows the battery performance and its recharge modalities, demonstrating the capability of the solar array to recharge them and provide the power to the other loads.

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Figure 10-5: Battery voltage and current (12:00 – 24:00 hrs Orbit)

Figure 10-6: Battery DOD (12:00 – 24:00 hrs Orbit)

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9:00 – 21:00 Hr Orbit In this case, just the battery performance results are showed because they already give a clear idea that the chosen configuration is suitable also for this mission, thanks to the contribution of the 5th panel on the satellite top.

Figure 10-7: Battery voltage and current (09:00-21:00 hrs orbit)

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Figure 10-8: Battery DOD (09:00-21:00 hrs orbit) 10.4.3 Mass/Power Budget

Table 10-3: Mass/power budget 10.4.4 Capabilities Provided to a Payload The power allocated for payload is 5 W (continous) or 10 W (with 50% duty cycle). Note that this does not meet the original study assumptions, but was necessary in order to bring the power demand down to an achievable level. 10.5 Scenario Study Case 2 (LEO With Propulsion) The addition of onboard propulsion is assumed to have no effect on the power budget, as manoeuvres will be scheduled to have no tangible effect on nominal operations.

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10.6 Scenario Study Case 3: GTO Mission 10.6.1 Design Drivers The GTO mission is characterised by a power demand of about 28 W in sunlight and 34 W in eclipse, with the maximum eclipse time being 2.2 hours. Three solar panels are required, one body mounted and the other two deployed as wings as illustrated in the Configuration chapter. The attitude is sun pointing, allowing for a constant illumination of the solar arrays during sunlight, but with different temperature profiles for the wings with respect to the body-mounted one. The table below reports the applicable power budget.

Table 10-4: GTO Power budget 10.6.2 Module Selection (Justification) The modules for this mission are: • 3 solar panels: 1 body mounted, 2 deployable • 3 MPPT solar array converters • 6 Battery modules. The following simulation results justify the need of the chosen configuration of the solar panels in order to provide the correct energy to the platform and the battery recharge.

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Figure 10-9: GTO Mission Energy Budget A further simulation over several orbits shows the battery performance and its recharge modalities, demonstrating the capability of the solar array to recharge them and provide the power to the other loads.

Figure 10-10: GTO Battery voltage and current

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Figure 10-11: GTO Battery DOD 10.6.3 Mass/Power Budget

Table 10-5: GTO Mass/Power budget 10.6.4 Capabilities Provided to a Payload As mentioned above, the power allocated for payload is 10 W (continuous).

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11 AOCS 11.1 Requirements and Design Drivers The requirements for this Nanosatellite platform are defined to enable a wide variety of missions to be implemented with the same platform. For this reason, the AOCS system is designed to be modular. These modules include the sensors, actuators and control software and are designed to enable the Nanosatellite to achieve the most stringent mission conditions/requirements that are foreseen for this platform. Any subset of these modules can be selected for each mission to achieve the specific requirements of each mission. 11.1.1 Attitude In order to allow the Nanosatellite platform to perform the types of missions that can currently be achieved with larger platforms, the following attitude motion requirements were defined: • Three axis stabilized, with good/high pointing knowledge and capability • Static Pointing, either inertial or celestial body pointing during normal operation mode • Achieve 45 to 90 degree slew manoeuvres within 300 to 400 seconds • Allow for an optional Spinning Platform up to 3 rpm. 11.1.2 Pointing Requirements The following requirements were derived from the systems level. They represent a compromise between the high performance required for a variety of current missions with large platforms and the maximum foreseen performance for the Nanosat mission, employing an extrapolation of the most accurate sensors and actuators foreseen for the mission.

Performance APE [arcmin] RPE [arcsec]/100s Description Minimum 6 60 Basic performance Maximum: 1 10 Best performance Table 11-1: Pointing Requirements 11.1.3 Orbital Scenario Compatibility • Platform compatible with LEO, GTO, GEO and Inerplanetary transfer orbits • Allow for optional Orbital Maintenance • Allow for De-Orbiting at EOL. 11.1.4 Target Budgets From the systems level, the mass and power allocations were derived by scaling from a larger platform, and are as shown below. Nevertheless, the target is to achieve the lowest possible mass and power consumption. • Target mass: less than 1.2 kg • Target Power Consumption: less than 2 W average.

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11.1.5 Assumptions The following assumptions about the spacecraft were taken into account for the design of the AOCS subsystem.

Assumptions Thruster Arm [m] 0.15 Satellite Mass [kg] 20 CG offset of CV [cm] 3 Recidual Diplole [Am^2] 0.05 Drag Coef. 2.1 Reflectance 0.2 Max Side [cm] 60 Min Side 30 Magnetorquer Dipole [Am^2] 0.07 wheel Capacity [Nms] 0.11 Table 11-2: Assumed Spacecraft Characteristics 11.2 Sub-System Description The AOCS system is composed of the classical components including sensors, actuators, determination algorithms and control laws. However, for this platform, each of the hardware components are encapsulated into modules containing a standard common interface. The result is that the modules can be plugged into the system as needed by each mission, allowing to optimize power, mass and performance as required. The modules include sensors, actuators and a modularized software and are listed below: • Digital Sun Sensor – Sensor, provides two axis low performance determination • Star Tracker – Sensor, provides three axis high performance attitude determination • Coarse Three axis Rate Sensor (Gyro) package – Sensor, provides attitude rate • Three axis Magnetometer – Sensor, provides two axis attitude determination • Navigation Camera – Sensor, provides position information • GNSS - Sensor, provides position information on a low earth orbit • Three axis Reaction Wheel package – Actuator, provides high precision torques • Three axis Magnetorquer Package – Actuator, provides torques to de-saturate the wheels • Modularized AOCS Software – Implements the determination algorithms and control laws. Compatible with any number and combination of sensor and actuator modules. Depending on each mission, different combinations of several of these modules can be selected to achieve the mission requirements. 11.2.1 Interfaces A key requirement for modularity is that the interface of the different modules be compatible. This allows for a “Plug & Play” modularity of the sensors and actuators. In addition, the interfaces for the modules that are directly exploitable by other missions, such as is the case for all the sensors, were selected to be compatible with most current space platforms. This allows for a more portable set of modules that can be employed in larger platforms. For non-portable modules, since compatibility with current platforms is not an issue, the interface was selected to be a simpler Transducer Bus with reduced mass and power.

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The following table summarizes the interfaces for the different hardware modules:

REF # MODULE INTERFACE

1 DIGITAL SUN SENSOR SPW

2 STAR TRACKER SPW

3 COARSE THREE AXIS RATE SENSOR (GYRO) PACKAGE SPW

4 THREE AXIS MAGNETOMETER SPW

5 NAVIGATION CAMERA SPW

6 GNSS SPW

7 THREE AXIS REACTION WHEEL PACKAGE TB

8 THREE AXIS MAGNETORQUER PACKAGE TB Table 11-3: Module Interface Definition (SpW: Space Wire interface; TB: Transducer Bus interface.) 11.3 Module Descriptions 11.3.1 Module 1 (Digital Sun Sensor) 11.3.1.1 Description and Overview The digital sun sensor provides attitude determination in two axes. It is based on an imager detecting light rays from the sun passing through a pinhole filter. The imager sensor, processor, and interface electronics are placed on the same chip. This sensor & interface “on a chip” reduces, mass and power consumption. The imager sensor required for this module is already available on the market. In order to realize this design, the main technological development needed is the integration on a single chip of the sensor and interface electronic. A Technology Research Program (TRP) study has confirmed the feasibility of the “Sensor on a Chip” technology and has yielded a prototype sun sensor, shown in Figure 11-1 below. Currently, coarse analogue sensors of similar size, mass and power already available but low RAD tolerance (Aeroastro, Sinclair interplanetary, TNO).

Figure 11-1: Digital Sun Sensor Prototype

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11.3.1.2 Key characteristics

Digital Sun Sensor Mass [kg] 0.060 Peak Power [W] 0.21 Absolute accuracy (3 σ) [Deg] 0.02 FoV [deg] 128 Update rate [Hz] 10 Max Angular Rate [rpm] 16 Dimensions [mm] 30x30x25 Table 11-4: Sun Sensor Characteristics 11.3.1.3 Scalability The performance of this sensor is not scalable. However, multiple sensors could increase the overall field of view, which can increase the flexibility in sensor placement. In addition, multiple sensors also adds redundancy to the system. 11.3.1.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications to its hardware. This module is powered by 5V and therefore would impose a requirement on any host platform to be able to provide the appropriate voltage supply. 11.3.2 Module 2 (Star Tracker) 11.3.2.1 Description and Overview The star tracker provides attitude determination in three axis by observing some of the stars in the sky and comparing the images to a star catalogue. This sensor is composed of an imager with optics and a processor. The imager sensor, processor, and interface electronics are placed on the same chip. This sensor & interface “on a chip” reduces, mass and power consumption. The imager sensors currently available on the market from Aeroastro are 500 – 600g and around 2 W, but these have a poor performance and are not Rad hard. (Typical is 2.5 kg with 9 W for state of the art STR). In order to realize this design, the main technological developments needed are the integration on a single chip of the sensor and interface electronic as well as the development of lower power and mass imager sensors. A Technology Research Program (TRP) study confirmed the feasibility of this target design and yielded the prototype star tracker casing shown in Figure 11-2 below.

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Figure 11-2: Star Tracker Casing Prototype. 11.3.2.2 Key Characteristics

Star Tracker Mass [kg] 0.175 Peak Power [W] 0.72 Accuracy (1 σ) [arcsec] 15 Noise (1 σ) [arcsec] 7 SEA [deg] 40 Update rate [Hz] 4 Dimensions [mm] 42x37x83 Table 11-5: Star Tracker Characteristics 11.3.2.3 Scalability The performance of this sensor is not directly scalable. It is mainly determined by the optics and therefore partial re-design would be required. Moreover, multiple sensors could increase the overall field of view, which can increase the flexibility in sensor placement. In addition, multiple sensors also adds redundancy to the system. 11.3.2.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications. 11.3.3 Module 3 (Coarse Three Axis Rate Sensor (Gyro) Package) 11.3.3.1 Description and Overview The coarse three axis gyro package provides body rate information in three axis. Combining the single axis gyros into a three axis package reduces mass, volume & power. Most missions employ 3-axis rate sensors, and therefore this module is optimized for most missions. The baseline design of this sensor uses the same detector as those available in current MEMS gyros (700g). The electronics for all three sensors can all be placed on a single mixed signal ASIC, bonded to the sensors in order to further reduce mass, volume and power. The single axis rate sensors are already available on the market. In order to realize this design, the main

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 146 of 267 technological development needed is the integration on a single mixed signal ASIC of the sensor electronics and interface electronics. 11.3.3.2 Key Characteristics

Coarse Three Axis Gyro Mass [kg] 0.106 Peak Power [W] 0.33 Bias [deg/hr] 10 Dias Drift [deg/hr/hr] 5 Noise [arcsec/sec] 17 Random Walk [deg/rt(hr)] 0.5 Update rate [Hz] 10 Dimensions [mm] 30x35x25 Table 11-6: Three Axis Rate Sensor Characteristics 11.3.3.3 Scalability The performance of this sensor is not scalable. Multiple sensors only add redundancy to the system. 11.3.3.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications. 11.3.4 Module 4 (Three axis Magnetometer) 11.3.4.1 Description and Overview The three axis magnetometer provides two axis attitude information by measuring the strength and direction of the magnetic field of the celestial body the spacecraft is orbiting and comparing it to an internal model of this magnetic field. The magnetic sensors, based on AMR (Anhysteretic Remanent Magnetization) detection, are already available off the shelf in CMOS technology. Incorporation of rest of the fairly simple electronics & interface on the same chip will reduce mass, volume and power consumption. In order to realize this design, the main technological development needed is the integration on a single chip of the sensor, electronics and interface.

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11.3.4.2 Key Characteristics

Three axis Magnetometer Mass [kg] 0.052 Peak Power [W] 0.22 Accuracy in [nT] 100 Update rate [Hz] 10 Dimensions [mm] 30x30x30 Table 11-7: Three Axis Magnetormeter Characteristics 11.3.4.3 Scalability The performance of this sensor is not scalable. Multiple sensors only add redundancy to the system. 11.3.4.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications. 11.3.5 Module 5 (Navigation Camera) 11.3.5.1 Description and Overview The navigation camera provides information about the position of the spacecraft by observing the stars and comparing the images to an internal catalogue. The navigation Camera hardware identical to that of the Star tracker, however the processing software is different. Please refer to the star tracker section for more information. 11.3.5.2 Key Characteristics

Navigation Camera Mass [kg] 0.175 Peak Power [W] 0.75 Update rate [Hz] 2 Dimensions [mm] 42x37x83 Table 11-8: Navigation Camera Characteristics

11.3.5.3 Scalability The performance of this sensor is not scalable. However, multiple sensors could increase the overall field of view, which can increase the flexibility in sensor placement. In addition, multiple sensors also adds redundancy to the system.

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11.3.5.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications. 11.3.6 Module 6 (GNSS) 11.3.6.1 Description and Overview The GNSS (Global Navigation Satellite System) provides position and velocity information about the spacecraft. The baseline GNSS design is a GPS. Off the shelf GPS with similar characteristics are already available on the market. The design is an extrapolation of what the products in the market will be like at the time of the Nanosatellite development. 11.3.6.2 Key Characteristics

GNSS Mass [kg] 0.175 Peak Power [W] 0.80 Time to first fix [s] 120 Position accuracy [m] 5 Update rate [Hz] 2 Dimensions [mm] 50x50x40 Table 11-9: GNSS Characteristics 11.3.6.3 Scalability The performance of this sensor is not scalable. Multiple sensors only add redundancy to the system. 11.3.6.4 Multi mission suitability This module is reusable in other classes of missions which require similar performance. The standard SpW interface makes the module reusable in other systems with no modifications. 11.3.7 Module 7 (Three Axis Reaction Wheel Package) 11.3.7.1 Description and Overview The reaction wheels provide control torques through changes in the rotational momentum of the wheels. Three wheels were packaged into a three axis package due to the fact all missions will require three wheels for 3-axis stabilization within the required performance, since the required performance can not be achieved with the other actuator modules provided. This package reduces mass and power on the interfacing electronics, connectors and wires. A number of miniature wheels are already available on the market. However, the Nanosat mission will require a custom wheel for inertia and torque combination, to be radiation hard and to possess the correct interface. The development of the reaction wheels entails a scaling of off- the-shelf components, which does not present any mayor technical challenge. A figure of the envisioned three axis reaction wheel package is shown in Figure 11-3 below.

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Figure 11-3: Three axis reaction wheel package 11.3.7.2 Key Characteristics

Three Reaction Wheel Package Mass [kg] 0.462 Power [W] (Peak | Ave) 9 | 2.85 Torque [mNm] 0.7 Mom storage [Nms] 0.11 Update rate [Hz] 10 Dimensions [mm] 100x100x100 Table 11-10: Three Axis Reaction Wheel Package Characteristics 11.3.7.3 Scalability Employing multiple units will multiply all the characteristics and performance by the number of units. 11.3.7.4 Multi mission suitability This module is sized for a nano-spacecraft. Scaling for other missions entails a complete redesign of the module since some of the components, such as the motor and electronics are not directly scalable.

11.3.8 Module 8 (Three Axis Magnetorquer Package) 11.3.8.1 Description and Overview The three axis Magnetorquer package provides two axis control torques around the local magnetic field vector via a generated magnetic field. They are employed mainly for reducing tip- off rates, specially during safe mode, and for reaction wheel off-loading. The Magnetorquers were packaged into three axis in order to reduce mass and power consumption in the interfacing

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 150 of 267 electronics. In addition, since Magnetorquers are most commonly used in 3-axis, packaging them into a 3 axis module does not present any foreseeable drawbacks. Similar miniature MTQ already exist, however the driver and interface electronics must be developed and integrated to the torquers, which presents a small challenge. Figure 11-4, below shows the envisioned configuration of the three axis Magnetorquer package.

Figure 11-4: Three axis Magnetorquer package 11.3.8.2 Key Characteristics

Three Axis Magnetorquer Power [W] (Peak | Ave) 3 | 2 Mass [kg] 0.101 Dipole [Am2] 0.07 Update rate [Hz] 10 Dimensions [mm] 90x90x90 Table 11-11: Three Axis Magnetorquer Characteristics 11.3.8.3 Scalability Employing multiple units will multiply all the characteristics and performance by the number of units. 11.3.8.4 Multi mission suitability This module is sized for a nano-spacecraft. Scaling for other missions entails a complete redesign of the module since some of the components, such as the coils and electronics are not directly scalable. In addition, Magnetorquers can only be employed in missions where the spacecraft is immersed in a magnetic field, making them only useful for low earth orbit missions or missions around other celestial bodies with magnetic fields of similar or higher strengths than that of the earth.

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11.3.9 Module 9 (Modular AOCS Software) 11.3.9.1 Description and Overview In order to reduce the recurring costs and the time associated with the development and AIT of the software with the AOCS hardware, the AOCS software module will be made compatible with and re-usable for multiple missions. This will be achieved by modularizing the software to include recurring blocks that will be utilized for all missions. Also, a pool of mode structures and control laws will be available to be selected by any given mission. A configuration data file, custom for every mission, will set all the relevant values for the specific mission and select which mode structures, mode transitions and control laws will be utilized. In addition, if any given mission requires custom mode structures or control laws that are not currently available in the existing pool, only those modules will need to be developed, which will also be available for subsequent missions, reducing the recurrent development for every mission. This architecture also allows for the development of a tool to rapidly re-configure the AOCS software, by generating the custom configuration data files, which further reduces the recurrent cost of AOCS software development for every mission and makes the AOCS even more flexible. Only a conceptual approach has been developed and a TRP study needs to corroborate its feasibility. However, Figure 11-5 below shows a concept block diagram of the software modules and their interactions.

AOCS Software Configuration Data Mode Structures

Mode 1: Mode Structure 3 Axis Pointing De-Spin Spinning Control law Parameters Estimator 2 Axis Pointing Parameters Slew 6 DOF 1 Axis Rate Mode 2: ….

Control Laws Mode Controller PID H- External Kalman Quaternion

External OBDH Command and TM Routines Control

Estimators

FDIR Checks Legend Profile Generators

Recurrent for every mission

Torque and Delta-V Equipment Checks Equipment Status Customizable Realiser and Calibrations

Custom blocks can be added

Equipment PnP Interface

Figure 11-5: Modularized AOCS Software block diagram

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In addition, it is foreseen that the software will have the size and processing power requirements listed in Table 11-12 below. These were extrapolated from current missions, as the AOCS software will perform similar operations to obtain similar performance.

AOCS Routines Min Typical

MIPS 6.75 Scale According to Cycle RAM [Mb] 1 2 NVM (Program) [Mb] 2 2 Options for 10, 2 and 1 Cycle [Hz] 4 Depending on application Table 11-12: AOCS Software computing requirements 11.3.9.2 Key Characteristics Modular multi-mission software. Customizable data files for each mission selecting modules and defining parameters. Custom Modes and/or control laws can be added if required by specific mission. Drawbacks: the flexibility of the software means that there will be large overhead and unused code. 11.3.9.3 Scalability Not applicable. 11.3.9.4 Multi mission suitability This module is reusable for any Nanosat mission with minimal development. Only configuration data files and any specific custom modes and control laws required by the mission need to be re- developed. In addition, if different sensors or actuators than those foreseen for this class of missions are required, the interfaces will need to be developed. 11.4 Scenario Study Case 1 (LEO – w/o Propulsion) 11.4.1 Design Drivers For this scenario, the design of the AOCS must achieve the required pointing performance while managing the reaction wheel momentum using Magnetorquers, since a propulsion system is not available. 11.4.2 Module Selection (Justification) The following list shows the modules selected for this scenario: • Digital Sun Sensor – Provides an attitude estimate during star tracker blackouts and in safe mode, can be used along with the magnetometer for a 3-axis attitude estimate • Star tracker – Provides the high accuracy attitude determination • Coarse three axis gyro – Used in conjunction with the star tracker for body rates estimation through a gyro-stellar estimation. Also used for de-spin and safe mode along with the digital sun sensors.

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• Three axis Magnetometer – Provides magnetic field information required for operation of the Magnetorquers. Can also provide an attitude estimate during star tracker blackouts and in safe mode, can be used with the sun sensors for a 3-axis attitude estimate • GNSS – Provides position knowledge • Three axis Reaction Wheel Package – Provides the fine control torques needed for fine pointing performance and high slew manoeuvres • Three axis Magnetorquer – Provides coarse torques for wheel momentum off-loading. 11.4.3 Mass/Power Budget

LEO Scenario Modules LEO Scenario Budgets

Table 11-13: LEO Scenario Budgets with NO Propulsion 11.4.4 Capabilities Provided to a Payload The following table describes the pointing performance achieved by the AOCS in this configuration:

Performance APE [arcmin] RPE [arcsec]/100s Determination 0.4 4 Pointing (Control) 1 10 Slewing 90 deg in < 100 sec

Table 11-14: LEO Scenario Performance with NO Propulsion These performance estimates are based on the best available sensor and actuator performance. In this case these are the star tracker and reaction wheels. The disturbance torques experienced by the spacecraft in this scenario, assuming the spacecraft characteristics shown in Table 11-2, are shown in Figure 11-6 below.

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Total Disturbance Torques -6 x 10 Worst Case in an 800km LEO - Max Torque 2.96e-006 Nm 3

2.5

1.5

0.5

0 Disturbance Torques [Nm] -0.5

-1

-1.5 0 0.5 1 1.5 2 2.5 3 3.5 4 Time [Orbits] (6052 seconds x Orbit) Figure 11-6: Disturbance Torques in the LEO Scenario The momentum accumulated in the wheels to reject the above disturbances is shown in Table 11-15 below. Off-loading the wheels using the Magnetorquers would require dumping, on average, for 92% of each orbit. During this time, on average, only 2/3 of the Magnetorquer power would be used since torque can only be produced perpendicular to the local magnetic field, which corresponds to one effective Magnetorquer axis at all times. Wheel off-loading LEO 800 km Frequency Source per Orbit Unit [Nms] Total / year Momentum increase due to Disturbances 1.00 0.006 31.27 Slew 45 deg 2.00 0.00006 0.63 Total Momentum 31.9 Magnetorquer Dump Torquer Dump (2/3 average) 100% 0.007 34.74 Total Dump will require 92% Table 11-15: Wheel Momentum Accumulation – LEO scenario According to this momentum accumulation rate, if no dumping is performed, the wheels will saturate within 14 orbits, as shown in Table 11-16 below. Wheel Saturation Orbit Type Momentum [Nms] Saturation Cyclyc + Slew Per orbit [Orbits] LEO 800km 0.024 0.006 14 Table 11-16: Wheel Saturation for the LEO Scenario

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11.5 Scenario Study Case 2 (LEO with Propulsion) 11.5.1 Design Drivers In this case, the propulsion system is a 1 DOF used for orbital maintenance, and the AOCS is identical as the above case. In the case where a 3-DOF propulsion system is employed, the design of the AOCS must achieve the required pointing performance while managing the reaction wheel momentum using the propulsion system. In this case, Magnetorquers would not be required. Since this would also be an interesting scenario, the resulting required ∆v for reaction wheel off-loading in this case is also shown in this section. 11.5.2 Module Selection (Justification) The modules used for this case are identical to those for the LEO case with no propulsion above. 11.5.3 Mass/Power Budget The budgets for this case are identical to those for the LEO case with no propulsion above. 11.5.4 Capabilities Provided to a Payload The capabilities provided to a payload are identical as the LEO case with no propulsion as discussed before. However, in the event that the reaction wheels are off-loaded using a propulsion system, 11 m/s of ∆v per year would be required, as shown in Table 11-17 below. The wheels would again saturate every 14 orbits, as discussed in the previous case, which entails that a propulsion manoeuvre would be required as a minimum every 14 orbits. Wheel off-loading LEO 800 km Frequency Source per Orbit Unit [Nms] Total / year Momentum increase due to Disturbances 1.00 0.006 31.27 Slew 45 deg 2.00 0.00006 0.63 Total Momentum 31.9 Propulsion System Dump (MIB* < 7 mNs / Thrust** > 1 mN ) Total Impulse [Ns] 213 Total Delta-V [m/s] 11 Table 11-17: Wheel Momentum off-loading using propulsion – LEO scenario 11.6 Scenario Study Case 3 (GTO) 11.6.1 Design Drivers For this scenario, the design of the AOCS must achieve the required pointing performance while managing the reaction wheel momentum using a combination of the propulsion system and the Magnetorquers. This is due to the fact that the Magnetorquers are only effective close to the perigee, since the rest of the orbit lies at distances from the earth where the magnetic field is too weak. In addition, it is not possible to use a GPS for GNSS due to the fact that for most of the orbit the spacecraft lies beyond the GPS satellites altitude, and therefore GPS is not available. The magnetometer is also not effective throughout the majority of the orbit due to its sensitivity and the fact that the magnetic field of the earth falls with the cube of the distance. Instead, an on-

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 156 of 267 board magnetic field model propagator is used along with the spacecraft attitude to determine the direction of the local magnetic field. 11.6.2 Module Selection (Justification) The following list shows the modules selected for this scenario: • Digital Sun Sensor – Provides an attitude estimate during star tracker blackouts and in safe mode, can be used along with the magnetometer for a 3-axis attitude estimate • Star tracker – Provides the high accuracy attitude determination • Coarse three axis gyro – Used in conjunction with the star tracker for body rate estimation through a gyro-stellar estimation Also used for de-spin and safe mode along with the digital sun sensors • Three axis Reaction Wheel Package – Provides the fine control torques needed for fine pointing performance and high slew manoeuvres • Three axis Magnetorquer – Provides coarse torques for wheel momentum off-loading, a magnetometer is not needed since it is assumed that an on-board magnetic field model propagator is used along with the spacecraft attitude to determine the direction of the local magnetic field • Propulsion System – This module is discussed in the propulsion system modules and is not included as an AOCS module. 11.6.3 Mass/Power Budget GTO Scenario Modules GTO Scenario Budgets

Table 11-18: GTO Scenario Budgets 11.6.4 Capabilities Provided to a Payload The following table describes the pointing performance achieved by the AOCS in this configuration:

Performance APE [arcmin] RPE [arcsec]/100s Determination 0.4 4 Pointing 1 10 Slewing 90 deg in < 100 sec

Table 11-19: GTO Scenario Performance These performance estimates are based on the best available sensor and actuator performance. In this case these are the star tracker and reaction wheels.

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The disturbance torques experienced by the spacecraft in this scenario, assuming the spacecraft characteristics shown in Table 11-2, are shown in Figure 11-7 below. Total Disturbance Torque -6 x 10 Best Case Orbit in an 600km - 35000km GTO 4

3.5

2.5

1.5

1 Disturbance Torque[Nm]

0.5

-0.5 0 0.5 1 1.5 2 2.5 3 3.5 4 Time [Orbits] (37415 seconds x Orbit)

Figure 11-7: Disturbance Torques in the GTO Scenario The momentum accumulated in the wheels to reject the above disturbances is shown in Table 11-20 below. Off-loading the wheels using the Magnetorquers dumping during 100% of each orbit would only dump 44% of the momentum. Note that the Magnetorquers are not effective for more than 2/3 of the orbit and therefore can be turned off during that time. Again, during the Magnetorquer on-time, on average, only 2/3 of the Magnetorquer power would be used since torque can only be produced perpendicular to the local magnetic field, which corresponds to one effective Magnetorquer axis at all times. The remaining momentum must be dumped using the propulsion system, which would require 1.2 m/s of ∆v per year. Off-loading the wheels using only the propulsion system would require 3 m/s of ∆v per year. Wheel off-loading GTO 600 km - 35000 km Frequency Source per Orbit Unit [Nms] Total / year Momentum increase due to Disturbances 1.00 0.009 7.59 Slew 45 deg 2.00 0.00006 0.10 Total Momentum 7.7 Magnetorquer Dump Torquer Dump (2/3 average) 100% 0.004 3.37 Total Dump will require 228% Torquers Effective only 20% of Orbit Propulsion System Dump (MIB* < 7 mNs / Thrust** > 1 mN ) Total Impulse [Ns] 51 3 Total Delta-V [m/s] *Based on 10% of wheel capacity.

**Based on offloading a full wheel in less than 10 minutes. Table 11-20: Wheel Momentum Accumulation – LEO scenario

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According to this momentum accumulation rate, the wheels will saturate within 10 orbits, as shown in Table 11-21 below, which entails that a propulsion manoeuvre will be required every less than 10 orbits. Wheel Saturation Orbit Type Momentum [Nms] Saturation Cyclyc + Slew Per orbit [Orbits] GTO 600 - 35000 km 0.02 0.009 10 Table 11-21: Wheel Saturation for the GTO Scenario 11.7 Telemetry & Data Requirements The following table summarize the telemetry requirements for the AOCS during different modes of operation. These apply to a spacecraft containing all the modules available for the AOCS and therefore constitute a worst case scenario.

AOCS Telemetry Rate [kbps] Comments

Typical @ 1Hz Cycle 1.8 Sub-Sampling High 7.2

Full (Troubleshooting) 18 Min (Safe Mode) 0.3 Table 11-22: AOCS Telemetry Requirements Summary

A summary of the data and interface requirements for all the modules available for the AOCS is shown below. These were the values assumed when calculating the telemetry budget summarized above.

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Comments interfaces 3 axis, number depends on final gyro configuration Three wheels, may only be one interface depending on configuration Three axes, may only be one interface depending on configuration Assuming groups of 3 thrusters per module, 12 thrusters with 4 Interface SpW SpW SpW SpW SpW TB TB TB stamp Time Stamp Stamp Time Accuracy [ms] 1 1 10 10 N/A 10 10 1 and duration [s] Latency Latency 0.05 0.05 0.05 0.05 N/A 0.05 0.05 Table 11-23: Modules Data Requirements Table 11-23: [Hz] Update Update 10 10 10 10 Clock GPS 10 10 10 No Used Typical 2 1 3 (1) 1 1 3 (1) 3 (1) 12 (4) AOCS Interfaces

Sun SensorSun Star Tracker Gyro Magnetometer GNSS (GPS) Reaction Wheel Magnetorquers Thrusters

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12 DATA HANDLING The current architecture of the distribution of signals, commands and data on European satellites is a “star” configuration using a data handling bus and remote units. Every single equipment is connected to a remote unit with a dedicated set of wires leading to a complex and heavy harness and an extended integration time to verify that each equipment has been connected correctly.

Figure 12-1: Schematic diagram of a classic 'star' system In addition to the trends for increased mission complexity, reliability, longevity and autonomy there is also a conflicting and real need for reduced mission costs. These trends are expected to continue into the next decade. A nanosat is the ideal opportunity to test all the new concepts that can bring real breakthrough towards the solution of these demands, but all aspects of the mission – from design, through unit procurement to AIT and operations - will need to be thoroughly examined and streamlined. More efficient architectures with greater standardisation and modularity while retaining the reliability are expected to be needed. It is now widely recognised that the absolute performance of avionic systems is dependent on an ability to transmit yet more data and in a more efficient way between systems. New applications need 'state-of-the-art' performance, not only for what concerns raw throughput levels, but mainly for ease of integration, management and flexibility of operations. The ever increasing requirement to have network enabled capabilities puts emphasis on compatibility too. Several network standards are already in use or potentially suitable for use in avionic systems. Some of the data network standards are specifically designed to meet the requirements of avionic systems (i.e. ARINC 429 and ARINC 629 and for specific space applications, MIL-STD-1553B

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 162 of 267 and prEN 3910). These, as well as standards designed for the even more demanding automotive/industrial systems (like CAN) will generally meet the environmental and real-time performance requirements of avionics systems. But for the specific case of a Nanosat it should be recognized that to achieve an ultimate reduction of the pin count and a higher level of integration of functions, a solution to remove ALL the point to point interfaces needs to be found and a very strict standardization, up to the SW API level, of the command, data and control interfaces is needed. It is the ability to concurrently design all parts of the S/C that makes this approach feasible.

Figure 12-2: Schematic diagram of a modern bus based system No single TM/TC architecture can be recommended as the best choice for all systems. Limiting the scope of choice to those architectures previously used or being considered by others goes a long way towards providing benefits of a single choice whilst maintaining the ability to choose a suitable standard commensurate with the system requirements. Although many of the COTS standards are able to provide this performance increase and probably meet the technical requirements of an avionics network, none are likely to be directly ‘fit for purpose’ without enhancement. Ultimately, the choice of architecture may be driven more by factors such as life cycle and component costs, implementation issues, maintainability, upgradeability, certification, etc., rather than by the primary characteristics or outright performance of the architecture.

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Advanced data systems (2015 – onwards) will be based on highly integrated ASICs, hosting multi-core processors, multi-cache levels, MMUs (already in study/development by TEC-ED), interconnected using command and control bus architectural elements as the classic Mil1553B, SpaceWire/Fibre, CAN/Flexray, Ethernet (AFDX). All communication protocols standardised E2E, based in many cases on standard Ips, ready to be put in ASICs or in high end FPGAs. Digital sensor busses are under study as well by TEC-ED since they allow to overcome the harness complexity generated by the 'slow' control and housekeeping needs and to cope with the additional autonomy requested to future systems. In a small platform it is of paramount importance to reduce the harness, not only for mass reasons. Use of standard interfaces simplifies AIV and guarantees unit’s reusability.

Figure 12-3: Digital sensor bus evolution In future systems, the traditional concept of determinism will be replaced by a more broader QoS guarantee, due to the extensive use of non-clock bound mechanisms as multi-level caches, scratchpads, deep pipelines, multi master busses, message prioritization.

Figure 12-4: Typical harness bundle

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The bulk of the harness in current space architectures lies in the multitude of point-to-point connections that (in star shape) need to be cabled between control units and peripherals. Several ESA studies have demonstrated that analogue connections only take between 50 and 80% of the overall harness mass. Bulky harness means placement constraints, more errors in manufacturing, difficult testing, resulting in increased costs during all AIV/T activities. 12.1 Requirements and Design Drivers 12.1.1 Budgets A real nanosat needs to have budget squeezed to the bare minimum. The provided envelopes for this nanosat CDF study were closer to a thin mini-sat than to a real nanosat. ASIC processes have already (at 180 nm, for example, remembering that current edge technology for silicon is below 45 nm, with COTS 32 nm chip expected early 2010) demonstrated the capability to produce 'spacecraft on a chip' data systems, that provide the same set of capabilities that were once given by a set of separate boxes.

Figure 12-5: Spacecraft on a Chip ASIC schematic A single chip provides: • LEON3-FT SPARC V8 processor macro including DSU, MMU and GRFPU-FT • 2x1553 BC/BM/RT modules, two CAN modules

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• TC decoder module, TM encoder module • Housekeeping module, CCSDS Time Management • 7 Spacewire modules • 4 UARTs modules (3 APB UART and 1 AHB UART) • memory controllers (SDRAM, SRAM, EEPROM) • AHB bus monitoring module • Two AHB status module • DMA Telemetry • Bridge & DMA between 2 AHB busses • A Switch Matrix module • IRQ controller, timers, GPIOs. Thus, the main design driver for the nanosat study was: • how many functions can we squeeze into a single (65 nm or beyond) ASIC keeping a hard power consumption ceiling of 2 W ?

OnceNV this Mass ASIC is defined, the rest of the architecture cascades off. NV Mass TMTC Memory Processi Processi Memory formatter ng ng Function Function Scratch OVP 12.2 Sub-System Description pad RAM Processi Processi Overview of the subsystem,ng how ngit is broken down intoSCCC modules and how modularity is achieved. This may requireFunction a diagram Function depending on the complexityEncoder Shaping Regulator Table giving overview of all the modules and theirI purpose Demodulator filter Mapping Q Shaping 12.2.1 Interfaces filter

DiagramSpW and description of Autonomousthe interfaces: DCDC TB PoCB router Thermal Converter a) with other subsystemsControl Master Master b) between modules within your subsystem

Figure 12-6: Nanosat Main DHS ASIC (enclosed in yellow) block schematic 12.3 Module Descriptions A high degree of integration between Power, DHS and Comms S/S has been proposed. This results in having an avionics core that manages all the “survival functions” of the S/ C and could be adapted to any S/C size. The items (not modules) that needs specific developments for nanosat are: • PCM:Power Conditioning Module (see Power)

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• DCM: Distribution/Command Module • DCB: Data/Cmd Bus (high speed) • PoCB: Power/Control Bus • TB: n too pessimistic. Transducer Bus (low speed).

Figure 12-7: Nanosat avionic schematic 12.3.1 Module 1 – Distribution & Control Module (DCM) A complex system that works is invariably found to have evolved from a simple system that works. Figure 12-5 shows the functionalities that are enclosed onto the DCM. We still believe that it is possible to have all the DCM, including Over Voltage protection, its dedicated DCDC converter and voltage regulator, in a single double euro PCB. 12.3.1.1 Description and Overview The DCM provides main processing functions thanks to its multi (4) core > 120 MHz processor. Use of multi core provides embedded TMR and/or application isolation (virtualization) if necessary. Such an architecture, if combined with fast data links, can provide on demand processing/non volatile memory for payloads. Multi-core CPUs are typically multiple CPU cores on the same die, connected to each other via a shared L2 or L3 cache, an on-die bus, or an on-die crossbar switch. All the CPU cores on the die share interconnect components with which to interface to other processors and the rest of the system. These components may include a front side bus interface, a memory controller to interface with RAM, a cache coherent link to other processors, and a non-coherent link to the I/O devices. The terms multi-core and MPU (which stands for Micro-Processor Unit) have come into general usage for a single die that contains multiple CPU cores.

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Minimum amount of NVRAM > 16 GiB (realistic usable target > 160 GiB). Due to its relatively simple structure and high demand for higher capacity, NAND Flash memory is the most aggressively scaled technology among electronic devices. The heavy competition among the top few manufacturers only adds to the aggression. Current projections show the technology to reach approximately 20 nm by around 2010. While the expected shrink timeline is a factor of two every three years per original version of Moore's law, this has recently been accelerated in the case of NAND flash to a factor of two every two years. In 2005, Toshiba and SanDisk developed a NAND flash chip capable of storing 1 Gb of data using Multi-level Cell (MLC) technology, capable of storing 2 bits of data per cell. In September 2005, Samsung Electronics announced that it had developed the world’s first 2 Gb chip. In March 2006, Samsung announced flash hard drives with a capacity of 4 Gb, essentially the same order of magnitude as smaller laptop hard drives, and in September 2006, Samsung announced an 8 Gb chip produced using a 40 nanometer manufacturing process. May 29, 2008 Intel Corporation and Micron Technology, Inc. introduced the industry's first sub- 40 nanometer (nm) NAND memory device, unveiling a 34nm 32 gigabit (Gb) multi-level cell chip. Thus with today's technology a multi chip module of 32 GiB is absolutely feasible. In 2015, provided that < 20 nm technologies are continuously screened for radiation effects and ad hoc controller solutions (in form of standard IP cores) are provided a 10-fold increase is even too pessimistic.

Figure 12-8: NAND Flash process accelerates Moore's law RAM 4 MiB (SRAM preferred). SRAM is more expensive and less dense, but faster and significantly less power hungry (especially idle) than DRAM. It is therefore used where either

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 168 of 267 bandwidth or low power, or both, are principal considerations. SRAM is also easier to control (interface to) and generally more truly random access than modern types of DRAM. Current fast SRAM chips have 4 Mb capacity, meaning that in 2015 an 16 MiB multi chip module would be easily feasible. On top of that, on chip scratchpad (~ 1 MiB) allows 0 ws average memory access (think it as an L3 cache, managed by SW...). The separation between the CPU and memory leads to the von Neumann bottleneck, the limited throughput (data transfer rate) between the CPU and memory compared to the amount of memory. In most modern computers, throughput is much smaller than the rate at which the CPU can work. This seriously limits the effective processing speed when the CPU is required to perform minimal processing on large amounts of data. The CPU is continuously forced to wait for needed data to be transferred to or from memory. Since CPU speed and memory size have increased much faster than the throughput between them, the bottleneck has become more of a problem. Use of multi level caches and on chip scratchpad, more than an increase of the CPU clock may give dramatic performances increases keeping power consumption low. Communication interfaces will be based on standard IP cores and will use 2.5 – 1.8 V operation (probably with 1.2 V core, with embedded LDO), at 65 nm technology or beyond, 5 M gates or above (excluding scratchpad). Compared with the SCOC3 in Figure 12-4, such a chip takes the integration of function beyond the data handling domain including power distribution and communication functions and a large scratchpad RAM. That said, its architecture is feasible and its prototyping is already possible using a large commercial FPGA (like Xilinx's VIRTEX5).

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Figure 12-9: Quad core Opteron AMD processor. Size of the die (45 nm) is approx 20 x 20 mm One of the biggest limits to this miniaturization is the number of available IO pins on a single package. The biggest available packages for space use are the bulky CCGA-524. An effort is needed to go towards BGA-type packages (even in flip-chip) that may guarantee the requested level of miniaturization.

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Figure 12-10: Scheme of a BGA package 12.3.1.2 Key characteristics 12.3.1.2.1 Power: Nanometer technology allows, thanks to the feature shrinking and lower gate voltages, higher clock rates with a minimum price in terms of power consumption. This, together with dynamic clock scaling techniques, has been already applied successfully in commercial technology for low power portable computing devices and also for mixed digital-RF comm chips (like the one used for wireless communications). So, while available computing power today (in the range of 70 MIPS - 35 MFLOPS @ 100 Hz) is deemed sufficient for a nanosat application, targeting a scalable clock device with advanced computing features (DSP functions, L2 cache, internal scratchpad) that may go up to 333 - 250 MHz (so with roughly three times the computing power) will allow greater flexibility and easier reuse of this costly ASIC, together with bigger performance margins for SW. DCU shall be in < 4 W total (including conversion), having an ASIC running at 250 MHz max (to limit the overall power consumption), plus memories and interface chips.1 So far is deemed not feasible to integrate large amount of RAM or NVRAM in ASIC, excluding the scratchpad. Each SpW interface at 20 Mbit costs ~ 250 mW, at 200 may reach 800 mW. Therefore, the use of high-speed interface will be limited to the bare minimum needed.

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Each TB interface costs ~ 5 mW @ 100 kbps (but with duty cycle ~ 0 …). Yet another possibility is the "clockless CPU" (asynchronous CPU). Unlike conventional processors, clockless processors have no central clock to coordinate the progress of data through the pipeline. Instead, stages of the CPU are coordinated using logic devices called "pipe line controls" or "FIFO sequencers." Basically, the pipeline controller clocks the next stage of logic when the existing stage is complete. In this way, a central clock is unnecessary. It might be easier to implement these high performance devices in asynchronous logic as opposed to clocked logic: • A complex System on a chip like this may have components that can run at different speeds in the clockless CPU. In a clocked CPU, no component can run faster than the clock rate. • In a clocked CPU, the clock can go no faster than the worst-case performance of the slowest stage. In a clockless CPU, when a stage finishes faster than normal, the next stage can immediately take the results rather than waiting for the next clock tick. A stage might finish faster than normal because of the particular data inputs (multiplication can be very fast if it is multiplying by 0 or 1), or because it is running at a higher voltage or lower temperature than normal. Asynchronous logic proponents believe these capabilities would have these benefits: • lower power dissipation for a given performance level • highest possible execution speeds. The biggest disadvantage of the clockless CPU is that most CPU design tools assume a clocked CPU (a synchronous circuit), so making a clockless CPU (designing an asynchronous circuit) involves modifying the design tools to handle clockless logic and doing extra testing to ensure the design avoids metastable problems. 12.3.1.2.2 Mass & form factor A single PCB shall hold all the DCU, for the moment let’s stick to the double euro format. So we try to be < 800 g, but magnetics may lift us up, even if now low profile, low mass magnetics are available and use of 8 V bus makes things much simpler. Interfacing: The DCU provides three types of combined interfaces 1. Power & Command 2. Power & Data (transducer Bus) 3. High Speed Routed Data link (Spacewire). Further discussion of the interfacing architecture to be found in Section 12.3.2 12.3.1.3 Scalability The module is intended to work as a standalone single chain controller. Cold redundancy can be foreseen simply duplicating the hardware.

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12.3.1.4 Multi mission suitability The same controller can act as payload/platform controller of larger missions. Of course the benefit, in > 1 ton satellites of a < 1 kg avionics are limited. 12.3.2 Module 2 – Remote Terminal Interface ASIC 12.3.2.1 Description and Overview Once the architecture of the core module is agreed, a standardized way to connect the RTs to the platform needs to be provided. There are three types of units connected to the DCM 1. Payloads or AOCS (complex) 2. AOCS (simple) 3. Slow control. The classification of these units is made thinking about a 3D matrix with: 1. Data throughput 2. Requested data latency 3. Power consumption. Units at the lowest corner of the matrix are slow control units that can use the TB only. Units with higher power consumption shall use the TB and power bus. Units with Higher data throughput or particular needs on data latency shall use Spacewire + power bus. In no cases is it foreseen to use all the three buses on a unit at the same time. This can lead to the development of a standard component that may hold all the external interfaces for all the units. Once the interfaces have been defined, prototyped and standardized, this component is very simple and may be built using common ASIC technology.

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+V IPM 3.3 V -V REG 1.8 V Cmd dec

SpW EDS router Handler

TB HK sensors slave (ADC/DAC/T)

Figure 12-11: Scheme of the general purpose RT interface ASIC For this component too, use of asynchronous (clockless) logic techniques would be very beneficial in keeping the power consumption to the bare minimum. This chip will handle the EDS (Electronic Data Sheet) of the attached unit, providing to the bus master all the needed information to connect, handle, FDIR the unit, in a quasi Plug'n'Play manner. It shall again be stressed that a higher level of interface standardization brings together systems that are easier to integrate at software and functional level. For the slow control part another simple mixed mode digital ASIC (that corresponds to the bottom right corner IP of the previous one) shall be developed. This ASIC will allow direct connection of all the slow control related systems (physical transducers, heaters, switches) to the transducer bus and will be directly powered via the bus.

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TB 8 x Analog IF input

Alar 10 bit ADC m

T 4 x 8 bit sens DAC

Figure 12-12: Scheme of the general purpose Slow control ASIC The resulting ASIC can be used as component inside units or standalone. For the nanosat application a solution to have it connected without use of bulky connectors to the bus, when used as simple transducer (the ASIC has an embedded T sensor) shall be envisaged. For example, the bare die (plus pigtails) can be embedded into kapton foils and used as a “stick on” transducer. The overall general architecture (that can then be also reused in all space embedded systems) is shown in the next figure:

Figure 12-13: Scheme of the general slow control architecture 12.3.2.2 Key Characteristics This element is just two standard components. Mass, power etc, shall be considered zero.

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12.4 Scenario Study Case 1 (LEO w/o Propulsion) For the DHS, all three modules listed above will be required in all cases (with the number of RT interface and slow control ASICS being directly accessible from the system design). 12.5 Scenario Study Case 2 (LEO with Propulsion) See Section 12.4 above. 12.6 Scenario Study Case 3 (GTO) See Section 12.4 above.

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13 TELECOMMUNICATIONS 13.1 Requirements and Design Drivers The requirements for the NanoSat communications sub-system are: • Up to 1 kg of mass is allowed for the whole sub-system, including antennas and cables • The communications sub-system must be optimized in terms of power consumption • In order to fit the design goals, the sub-system shall provide modularity and scalability. On one hand, modularity will allow reconfiguring the sub-system for different missions by simply changing the modules for other ones. On the other hand, the scalability will ensure that a given module covers a maximum number of different scenarios. 13.2 Sub-System Description A typical onboard communications sub-system is a set of modules that receive and send digital data to and from the DHU. It performs the adaptation between the digital information in the DHU and the radiofrequency signals received and sent by the antennas. It is usually made up of several units: a transmitter, a receiver, an SSPA and a Radio Frequency Distribution Unit and antennas. The total mass including the antennas is about 3kg. In order to reduce the mass the best solution is to reduce the number of modules.

Digital RF

Transmitter SSPA DHU Receiver RFDU Antennas

Figure 13-1: Example of a typical communications sub-system As shown in Figure 13-1, the sub-system can be divided in two parts depending on the type of signals they deal with, either digital or radiofrequency. In order to reduce mass and complexity of the cabling, the proposed solution for NanoSat is based on the division of the digital part and the analog part in two separated units, as shown in Figure 13-2:

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Digital RF

UP/DW

DHU converter Modem Antennas

TELECOMMANDS

TELEMETRY Figure 13-2: NanoSat communications sub-system The proposed solution takes advantage of the already existing digital resources of the Data Handling Unit to implement the digital part of the communications sub-system. Indeed, all the digital blocks shall be integrated within the DHU and shall share the same ASIC and structure, thus reducing the overall mass and cost. The outputs/inputs of the DHU are not digital signals anymore but a low frequency modulated RF signal. All the RF blocks will be implemented in a separate unit called up/down converter. This unit will act as the interface between the DHU modem and the antennas. Finally, two antennas placed on opposite sides of the spacecraft are used to provide an omnidirectionnal pattern. 13.2.1 Modularity and Scalability Only one model of the DHU Modem and the antenna is required, covering all possible scenarios. The sub-system modularity and scalability will be achieved by developing a set of UP/DW converters, each one providing a different level of RF output power. The appropriate UP/DW converter shall be chosen depending on the available onboard power. In total, four models providing 1, 2, 4 and 8 Watts shall be developed. The basic idea is that doubling the RF output power will double the amount of data that can be dumped to Earth but also the sub-system consumption. Thanks to this, a trade-off between consumed power and data budget can be done. 13.2.2 Interfaces Figure 13-3: shows the interfaces between communications modules and other sub-systems:

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Onboard Demodulator ADC AGC LNA 3dB Computer cos() LO Shaping I filter SCCC sin() Mapping DAC SSPA Encoder Q Shaping filter

DHU Modem UP/DW Converter Antennas Power Figure 13-3: NanoSat communications sub-system detailed diagram Two other sub-systems are interfaced with the communications subsystems. On one hand the Data Handling Unit shall have two SMA interfaces providing one input and one output for telecommands (uplink) and telemetry (downlink) respectively. A third data interface shall be used to control RF switch selecting the transmitting antenna. The interface between the UP/DW converter and the antennas shall be to SMA cables, one for each antenna. Finally the DHU modem and the UP/DW converter must be connected to the power sub-system. 13.2.3 Modules List Table 13-1 shows a list with all the modules that make up the baseline communication subsystem (note that the baseline communications subsystem is an S-band system):

Module name Mass Dimensions Pon Pstb DHU Modem + 100 g - + 0.1 W + 0.1 W S-band 1W UP/DW converter 900 g 100x100x40 2.5 W 1 W S-band 2W UP/DW converter 900 g 100x100x40 5 W 1 W S-band 4W UP/DW converter 900 g 100x100x40 10 W 1 W S-band 8W UP/DW converter 900 g 100x100x40 20 W 1 W S-band LGA 20 g 50x50x10 0 W 0 W Cables 50 g 0 0 W 0 W Table 13-1: Modules list (baseline) It has to be pointed out that the mass and power consumption of the DHU modem has to be considered in the DHU sub-system budgets. The list of the necessary equipment to implement the NanoSat communications sub-system is: • 1 DHU modem • 1 UP/DW converter (to be chosen among the 4 available models) • 2 Antennas • Set of cables.

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Table 13-2: shows a list with all the other modules that have been considered in the study but were not kept as the baseline. These modules will be presented and assessed in chapter 13.7 (Other Solutions and Trade-Offs):

Module name Mass Dimensions Pon Pstb X-band-EO 1W UP/DW converter 900 g 100x100x40 2.5 W 1 W X-band-EO 2W UP/DW converter 900 g 100x100x40 5 W 1 W X-band-EO 4W UP/DW converter 900 g 100x100x40 10 W 1 W X-band-EO 8W UP/DW converter 900 g 100x100x40 20 W 1 W X-band-SR 1W UP/DW converter 900 g 100x100x40 2.5 W 1 W X-band-SR 2W UP/DW converter 900 g 100x100x40 5 W 1 W X-band-SR 4W UP/DW converter 900 g 100x100x40 10 W 1 W X-band-SR 8W UP/DW converter 900 g 100x100x40 20 W 1 W X-band LGA 100 g 30x30x40 0 W 0 W Mobile phone based transponder 300 g 100x100x40 1 W 0.1 W Multifunctional distributed antennas 10 g 100x10x10 0 W 0 W Electronics for distributed antennas 150 g 100x30x15 2 W 0.1 W Table 13-2: Modules list (options) 13.3 Module Descriptions 13.3.1 Module 1 (DHU Modem) 13.3.1.1 Description and Overview This module is in fact included within the DHU. The DHU modem contains all the digital elements of the communications sub-system that shall be integrated in the DHU ASIC. Figure 13-4: depicts the block diagram of the DHU modem:

Onboard Demodulator ADC AGC Computer cos() Shaping I filter SCCC sin() Mapping DAC Encoder Q Shaping filter

Figure 13-4: DHU modem The upper chain contains the elements used by the uplink demodulator. The uplink modulation scheme shall be a typical NRZ/PSK/PM. The lower chain contains the elements used by the downlink modulator. The downlink modulation scheme shall be a SCCC (Serially Concatenated

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Convolutional Codes) (see RD[25]). In both chains the digital elements are represented with yellow boxes while a few RF elements (AGC, ADC, DAC and connectors) are represented with blue boxes. The input of the demodulator is an RF modulated signal at intermediate frequency (around 100 MHz). The SCCC modulator output is an RF modulated signal at variable intermediate frequency (between 100 MHz and 190 MHz). The downlink modulation scheme (SCCC) is currently being proposed in CCSDS and might be approved as the new standard for high data rate telemetry links in a few years. An SCCC laboratory demonstrator has already been developed and tested, confirming its good performance. However, it is still considered as an orange book and it has not been flown yet. On the ground segment side, an SCCC receiver shall also be developed. However, a high data rate SCCC receiver, based on a Cortex platform, is currently being developed and will be ready in two years. It seems feasible to expect that within 10 years, high and low data rate SCCC receivers will be available on the market. As far as the uplink demodulator is concerned, the NRZ/PSK/PM modulation has been used as a CCSDS telecommands standard for years and no special developments are foreseen. The integration of a modem in the DHU should be feasible without major issues. 13.3.1.2 Key characteristics The most important characteristic of DHU modem is the use of SCCC modulation in the downlink. This scheme provides variable bit rate at fixed symbol rate (constant channel bandwidth) and it also takes advantage of advanced coding techniques increasing the maximum data throughput. SCCC also offers different roll-off factors (0.2, 0.25, 0.30 and 0.35) in order to optimize the occupied bandwidth. The DHU modem shall provide variable downlink symbol rate (from 50 ksps to 5 Msps) so that it can be adapted to the mission requirements. In terms of implementation is has to be seen that all the digital elements will be implemented in the DHU ASIC, requiring about 100 kgates. Thanks to its implementation on an ASIC, the power consumption of the digital part of the modem shall be reduced by a factor 100. The estimated mass of the RF elements including the connectors is 100 g and their power consumption is 100mW (in both, transmission and reception modes). The DHU modem is based on a software radio scheme, where frequency conversion from baseband to intermediate frequency is performed in digital. This frequency conversion determines the necessary clock frequency in the DHU that has to be at least double of the output frequency. As a consequence, a minimum of 400 MHz clock is required in the DHU. The estimated power consumption of the DHU modem is 100mW continuously (the demodulator must be always kept on).

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13.3.1.3 Scalability Thanks to SCCC data rate flexibility is achieved by choosing the appropriate modulation and coding scheme (hereafter called ModCod) and keeping a constant symbol rate. Three different modes are available: • Constant Coding and modulation (CCM): Typical approach • Variable Coding and Modulation (VCM): ModCod changes during the pass. The ModCod sequence must be preset onboard via TC before the pass • Adaptive Coding and Modulation (ACM): ModCod dynamically changes during the pass. The ModCod is set via TC by the G/S depending on the received SNR. It requires a return link during transmission. Depending on the SNR, up to 27 ModCods are available. The most robust ModCod is a QPSK 2/5 with a spectrum efficiency of 0.75 bits per symbol and required Es/No = 0.5 dB. The less robust ModCod is a 64APSK 9/10, with a spectrum efficiency of 5.39 and a required Es/No = 19.1 dB. The total range of signal to noise ratio is thus 18 dB. In a link budget, this can be seen as 8.5 times of range increase (for example, from 3000 km to 25000 km). As an example, the following figure shows how SCCC in VCM mode works. When the distance between the S/C and the G/S increases the signal to noise ratio of the telemetry downlink decreases. The VCM allows adapting the bitrate to the signal to noise ratio, thus sending the maximum number of bits per symbol at any time:

15 1000 948 Variable Coding and Modulation (400ksps) 12 800 732

9 600 556 484

Es/No [dB] 6 416 400

344 344 Data Rate [kbps]

3 200 QPSK 16APSK 8PSK QPSK 3/7 3/5 5/8 2/3 0 0 4000 5000 6000 7000 8000 9000 10000 Range [km]

ES/No req Margin Data rate

Figure 13-5: SCCC in VCM example

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13.3.1.4 Multi mission suitability The variable output frequency allows choosing the downlink frequency within the 90 MHz band available in S-band, thus spreading all the NanoSat transmissions within the S-band. The variable symbol rate is essential in order to adapt the transmission to different mission scenarios with different ranges. 13.3.2 Module 2 (UP/DW Converter) 13.3.2.1 Description and Overview A simplified block diagram of the UP/DW converter is depicted in Figure 13-6:

LNA 3dB

SSPA

Figure 13-6: UP/DW converter The lower chain corresponds to the down converter which is used in the downlink transmission. In this chain, a mixer converts the intermediate frequency signal coming from the DHU to an S- band signal. The signal is then amplified by an SSPA. Finally, an RF switch controlled by the OBC allows selecting the transmitting antenna, depending on the satellite attitude. The upper chain corresponds to the up converter which is used in the uplink reception. The up converter starts with two hybrids that split the uplink and downlink signals on each antenna. The uplink signals from both antennas are then merged with a 3dB combiner. The uplink signal is then amplified by a Low Noise Amplifier. Finally, a mixer converts the signal from S-band to intermediate frequency. UP/DW converters modules are largely used in ground stations. In space, they are split in several parts and integrated with other modules. For example, the down converter part (LNA, mixer and filter) are included in the receiver. Thus the UP/DW converter is usually split in three different modules: receiver, SSPA and RFDU. The main advantage of the proposed configuration is that all the RF components are in the same box, thus saving the required mass. The main technology developments for the UP/DW converter development are: 1. Implement a class F SSPA in order to improve the power efficiency. New class SSPA is currently being investigated by many companies and agencies, but current flight tested equipment including COTS only offer about 25% of power efficiency.

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2. Implement the radiofrequency circuits using GaN instead of GaAs in order to increase the power efficiency. GaN technology is currently being investigated in several activities at ESA. 3. Finally a module integrating the up converter (mixer and SSPA), the down converter (LNA and mixer) and the RFDU (hybrids, 3dB combiner, and switch) has to be developed. In order to cope with different mission scenarios four different UP/DW converter models shall be developed, each one providing a different RF output power (1, 2, 3 and 8W). The design of the module in terms of mass and size remains the same, being the SSPA the only element in the module that changes. 13.3.2.2 Key Characteristics The key characteristic of this module is the power consumption, mainly required by the SSPA. Due to the nature of the radiofrequency waves’ attenuation with the distance, the downlink signal must be sufficiently amplified in order to reach the G/S receiver a minimum power level required to correctly demodulate the signal. The required RF output power increases with the square of the distance. This means that when distance doubles the necessary output power on the S/C must be multiplied by 4. In addition, the necessary output power also increases with the symbol rate, so that doubling the symbol rate requires doubling the output power. This means that the RF output power is actually a driver given by the communication requirements and not a variable parameter on the communication sub-system design. Nevertheless, big improvements can be done in terms of power efficiency. The UP/DW converter proposed for NanoSat shall make use of new class SSPAs, providing power efficiencies of at least 50%, instead of the current 25% efficiency. It will also make use of GaN technology instead of GaAs, which can improve even more the power efficiency. Table 13-3: shows the foreseen power consumption of the UP/DW converter for NanoSat (note that the four different modules are listed):

DC Power DC power (stb)

UP/DW 1W 2..5 W 1 W

UP/DW 2W 5 W 1 W

UP/DW 4W 10 W 1 W

UP/DW 8W 20 W 1 W

Table 13-3: UP/DW converter power consumption Total mass of the module is 900 g with a size of 100x100x40 mm. 13.3.2.3 Scalability The four different modules to be developed shall be based on the same structure and only the SSPA shall change. The size, mass and connectors are then the same for all modules. Thus, one module can be easily replaced by another depending on the mission requirements.

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13.3.2.4 Multi mission suitability The scalability of the module is achieved by developing 4 different modules each one providing a different RF output power level. The UP/DW converter shall be chosen depending on the mission requirements, namely the distance between the S/C and the G/S, the required symbol rate and the available onboard power. 13.3.3 Module 3 (S-band LGA) 13.3.3.1 Description and Overview Two isogain S-band antennas shall be placed on opposite sides of the spacecraft in order to provide an omnidirectional pattern. The antenna type is a patch one, since it is easy to accommodate due to its small size and weight. Figure 13-7: shows a photo of an SSTL S-band patch antenna, currently used in many ESA missions:

Figure 13-7: UP/DW converter The antenna proposed for NanoSat is a modified version of the SSTL one. In order to reduce its mass the metallic plate of the antenna shall be removed and the radiating element shall be assembled directly on the satellite structure. 13.3.3.2 Key Characteristics The radiation pattern of each antenna must cover 180 degrees with more than -3dB of gain. Total mass of the each antenna is estimated to 20 g with a size of 50x50x10 mm. Logically, the antenna does not consume any power. 13.3.3.3 Scalability Only one model of S-band LGA will be used. 13.3.3.4 Multi mission suitability The same S-band LGA can be used for all missions.

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13.4 Scenario Study Case 1 (LEO w/o Propulsion) 13.4.1 Design Drivers The LEO scenario drivers for the communications sub-system are: • The main goal is to provide the spacecraft with a high data rate link in order to cope with a payload that could generate big amounts of data. It is foreseen that, in its nominal configuration, the communications sub-system shall be able to transmit data at 2 Mbps • The second driver is the use of a small G/S antenna. Even though the use of big G/S antenna is not ruled out, the communications sub-system shall provide a high data rate link using a 3.4 m antenna • Due to the nature of the platform, the available onboard power may be limited. As a consequence the communications sub-system shall make a good use of the available power and maximize the ratio data transmitted vs. power consumed. 13.4.2 Module Selection The following modules are selected for the LEO study case: • 1 DHU modem • 1 S-band UP/DW converter 2W • 2 S-band LGA • Cables. The choice of the UP/DW converter is been based on the power budged of the LEO scenario. However, other models of UP/DW converters could be used depending on the available power. The data budget results that will be shown in this section can be directly scaled to another power. Indeed, using a 4W UP/DW converter would double the amount of transmitted data while using a 1W UP/DW converter would halve it. 13.4.3 Mass/Power Budget Table 13-4: shows the dimensions, masses and power consumption for both, sunlight and eclipse modes:

Sunlight Eclipse Duty Duty Module name Quantity Mass Dimensions Pon Pstb Pon Pstb cycle cycle DHU Modem 1 + 100 g - + 0.1 W + 0.1 W 15% + 0.1 W + 0.1 W 30% S-band UP/DW 1 900 g 100x100x40 5 W 1 W 15% 5 W 1 W 30% converter 2W S-band LGA 2 20 g 50x50x10 0 W 0 W 0% 0 W 0 W 0% Cables 1 50 g 0 0 W0 W0%0 W0 W0% TOTAL 990 g 5 W 1 W 0 W 5 W 1 W 0 W Table 13-4: Scenario Study Case 1 (LEO) mass and power budget

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Please, note that the total mass and power values do not take into account the DHU modem contribution. Indeed, the DHU modem mass and power consumption shall be taken into account in the DHU budgets. 13.4.4 Capabilities Provided to a Payload The transmission mode for this study case is the Constant Coding and Modulation one (CCM). The choice has been done taking into account that VCM mode only increases the data budget by 44%. Such increase is not worth the increase of the uplink commands, necessary to set the ModCod sequence before each pass. The chosen ModCod is a QPSK3/7 at 3 Msps, which offers an efficiency of 0.86 bits/symbol. The useful data rate is thus 2.58 Mbps. Table 13-5: shows the telemetry link budget computed for a 3 Msps transmission. The link margin in this case is 3.82 dB which is enough to ensure correct reception of data.

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Frequency GHz 2.245 Altitude km 800 Telemetry rate ksps 3000

Tx Power W 2.00 Tx losses dB 1 Tx antenna gain dB -3 Tx EIRP dBW -0.99

Elevation deg 10 Slant range km 2366 FSL dB 166.95 Atmospheric losses dB 0.5 Polarization mismatch dB 0.21

G/S antenna diameter m 3.4 G/S antenna G/T dB/K 12.04 G/S received power dBW -134.60

Rx S/No dBHz 70.99 Rx noise figure dB 1 Dem. Loses dBW 1 Es/No dB 4.22 Modcod 2 Mod QPSK Coding rate 3/7 Efficiency bits/symb 0.86 Es/No req. dB 0.40 Usefull data rate kbps 2580 Margin dB 3.82 Table 13-5: Scenario Study Case 1 (LEO) link budget

The total amount of transmitted data can be computed by simply multiplying the useful data rate by the G/S – S/C contact time. Considering Redu as the G/S and a minimum elevation angle of 10 degrees the mean contact time is 29.26 min/day. The maximum amount of data that can be transmitted is thus 2.58 Mbits/sec x 29.26 min/day = 4.53 Gbits/day. It has to be noted that in order to obtain the mean contact time all the passes has been taken into account. Ruling out short passes would reduce the data budget. 13.4.5 Scalability For the sake of completeness an assessment of the communications sub-systems scalability will be presented in this section.

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Two different variables will be assessed: the ground station antenna size and the RF output power. The results are shown in Table 13-6:

G/S antenna Ouput power Symbol rate Mod Data 1 W 1500 ksps QPSK 2.3 Gb/day 3.4 m 2 W 3000 ksps QPSK 4.5 Gb/day 1 W 5000 ksps 16APSK 23.2 Gb/day 15 m 2 W 5000 ksps 32APSK 30.7 Gb/day Table 13-6: Scenario Study Case 1 (LEO) scalability 13.5 Scenario Study Case 2 (LEO with Propulsion) As per Study Case 1. 13.6 Scenario Study Case 3 (GTO) 13.6.1 Design Drivers The GTO scenario drivers for the communications sub-system are: • The main goal is to provide the spacecraft with a high data rate link in order to cope with a payload generating a big amount of data • The second driver is the use of a small G/S antenna. Even though the use of big G/S antenna is not ruled out, the communications sub-system shall provide a high data rate link using a 3.4 m antenna • The main issue of this scenario is that the range distance between the spacecraft and the ground station can vary from 600 to 40000 km. It is expected that the payload will be active in the apogee, thus the transmission scheme must be optimised to operate in the proximity of the Earth, below 30000 km. 13.6.2 Module Selection The following modules are selected for the GTO study case: • 1 DHU modem • 1 S-band UP/DW converter 4W • 2 S-band LGA • Cables. The choice of the UP/DW converter is based on the power budget for this study case. However, other models of UP/DW converters can be also used depending on the available power. The data budget results that will be shown in this section can be directly scaled to another power. Indeed, using an 8W UP/DW converter would double the amount of transmitted data while using a 2W UP/DW converter would halve it. 13.6.3 Mass/Power Budget Table 13-7: shows the dimensions, masses and power consumption for both, sunlight and eclipse mode of the modules selected for this study case:

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Sunlight Eclipse Duty Duty Module name Quantity Mass Dimensions Pon Pstb Pon Pstb cycle cycle DHU Modem 1 + 100 g - + 0.1 W + 0.1 W 15% + 0.1 W + 0.1 W 30% S-band UP/DW 1 900 g 100x100x40 10 W 1 W 15% 10 W 1 W 30% converter 4W S-band LGA 2 20 g 50x50x10 0 W 0 W 0% 0 W 0 W 0% Cables 1 50 g 0 0 W0 W0%0 W0 W0% TOTAL 990 g 10 W 1 W 0 W 10 W 1 W 0 W Table 13-7: Scenario Study Case 1 (GTO) mass and power budget Please, note that the total mass and power values do not take into account the DHU modem contribution. Indeed the DHU modem mass and power consumption shall be taken into account in the DHU budgets. 13.6.4 Capabilities Provided to a Payload The transmission mode for this scenario is the Variable Coding and Modulation one (VCM). In this scenario the large range of distances between the G/S and the S/C requires the transmission scheme to be gradually adapted, maximising the bitrate as a function of the distance. The ModCod shall then vary by steps of 1000 km of distance, in order to obtain the maximum data rate with a link margin of at least 3 dB. Table 13-8: shows the telemetry link budget computed for a 400 ksps transmission. For the sake of simplicity only two ModCods are shown. They correspond to the link in the 3000 – 4000 km range and the link in the 10000 – 11000 km range. It is clear that thanks to VCM the link margin is always close to the strict minimum, thus sending the maximum amount of data when the distance between the S/C and the G/S is small.

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Frequency GHz 2.245 2.245 Telemetry rate ksps 400 400

Tx Power W 4.00 4.00 Tx losses dB 1 1 Tx antenna gain dB -3 -3 Tx EIRP dBW 2.02 2.02

Elevation deg 10 10 Slant range km 4000 11000 FSL dB 171.51 180.30 Atmospheric losses dB 0.5 0.5 Polarization mismatch dB 0.21 0.21

G/S antenna diameter m 3.4 3.4 G/S antenna G/T dB/K 12.04 12.04 G/S received power dBW -136.15 -144.94

Rx S/No dBHz 69.44 60.65 Rx noise figure dB 1 1 Dem. Loses dBW 1 1 Es/No dB 11.42 2.63 ModCod 13 1 Modulation 16APSK QPSK Coding rate 3/5 1/3 Efficiency bits/symb 2.37 0.71 Es/No req. dB 8.00 -0.50 Usefull data rate kbps 948 284 Margin dB 3.42 3.13 Table 13-8: Scenario Study Case 1 (GTO) link budget The total transmitted data can be computed by multiplying the useful data rate by the G/S – S/C contact time in each 1000 km section and then summing them all. For the contact time computation, a G/S in Malindi has been considered. Low latitude ground stations offer the advantage of providing contact at short distances. For this G/S, the window between 3000 and 11000 km has been allocated for telemetry downlink, with a minimum elevation angle of 10 degrees. The mean contact time is thus 36.8 min/day and the maximum amount data that can be transmitted with (using VCM) is 1.06 Gbits/day. It has to be noted that in order to obtain this mean contact time, all the passes has been taken into account. Ruling out short passes would reduce the data budget. 13.6.5 Scalability For the sake of completeness an assessment of the communications sub-systems scalability will be presented in this section.

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Two different variables are assessed, the size of the ground station antenna and the RF output power. The results are shown in Table 13-9:

Antenna Power Symbol rate Time on Data 4 W 400 ksps 37 min/day 1.1 Gb/day 3.4 m 8 W 800 ksps 37 min/day 2.1 Gb/day 4 W 4000 ksps 79 min/day 26.9 Gb/day 15 m 8 W 5000 ksps 79 min/day 48.9 Gb/day Table 13-9: Scenario Study Case 1 (GEO) scalability 13.7 Other Solutions and Trade-Offs In this section other solutions for the communications sub-system are analysed. 13.7.1 X-Band vs. S-Band The use of X-band instead of S-band for the uplink and the downlink can be an interesting option offering different advantages and disadvantages. Since the choice of the frequency only modifies the output/input frequency, its choice only determines the UP/DW converter and the antenna design. Table 13-10: presents the main differences between X-band and S-band:

S-band (2 GHz) X-band (8 GHz)

Bandwidth 6 MHz EO: 375 MHz SR: 10 MHz

Symbol rate 5 Msps EO: 312.5 Msps (SRRC 0.2) SR: 8.3 Mbps

Frequency Same frequency for Different frequencies for EO and SR. EO and SR.

Antenna mass 20g 100g – 200g

Antenna size Patch Helix 30x30x40mm 50x50x10mm

Antenna gain Low High (+9 dB for isoflux)

Omnidirectional Yes No

Antenna TRLo 6 2

Technology Coaxial Waveguide

Table 13-10: X-band vs. S-band 13.7.1.1 Frequency band The allocated spectrum for space-to-Earth S-band links covers from 2100 MHz to 2195 MHz. The allocated spectrum for space-to-Earth X-band links is split in two different bands depending on the mission type. For Earth Exploration Satellite Services the so called EESS band is allocated. This band covers from 8025 to 8400 MHz. For Space Research (SR) a different band from 8450 MHz to 8500 MHz is allocated. This has an impact on the UP/DW converter design.

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When S-band is used the whole set of UP/DW converters can be used for both types of mission. However, when X-band is used the UP/DW converter must be designed either for EESS or SR. Thus, in X-band two sets of 4 UP/DW converters with different output frequencies must be developed. Another solution for X-band would be to provide the UP/DW converter reconfigurable output frequency, but that would lead to an increase of the module complexity. Since scalability is a requirement for NanoSat the S-band is the most suitable option in this case. 13.7.1.2 Bandwidth (or symbol rate) The maximum occupied bandwidth for S-band is 6 MHz. This limits the symbol rate (channel symbols) to 5 Msps (considering a SRRC filtering with 0.2 roll-off factor). The maximum occupied bandwidth in X-band is 375 MHz for EESS and 10 MHz for SR. This limits the symbol rate to 312.5 Msps and 8.3 Msps respectively. It is clear than the use of X-band offers more bandwidth and thus a higher symbol rate can be used, especially for EESS. As a consequence, X-band offers more data rate when link conditions are favourable (high RF output power and/or large G/S antenna). However, in S-band the symbol rate is limited and the only way to increase the data rate is using high order modulations and coding schemes. Table 13-11 compares the achievable data rate in both cases:

S/No 80.2 dB 80.2 dB 80.2 dB Symbol rate 5 Msps 37 Msps 5 Msps Es/No 13.2 dB 4.5 dB 13.2 dB Modcod QPSK1/2 QPSK1/2 16APSK3/4 Efficiency 1.0 1.0 2.9 Data rate 5 Mbps 38 Mbps 15 Mbps Req. Es/No 1.5 dB 1.5 dB 10.2 dB Margin 11.7 dB 3.0 dB 3.0 dB Table 13-11: X-band vs. S-band data rate comparison Columns two and three shown how the spare margin in column one can be used to increase the data rate by either increasing the symbol rate and keeping the same modulation and coding scheme (X-band case) or keeping the same symbol rate but increasing the modulation and coding scheme (S-band case). It is clear that increasing the symbol rate offers more data rate than increasing the modulation and coding scheme and thus the X-band seems the more suitable option for high data rate downlinks. 13.7.1.3 Antennas Another important difference between S-band and X-band is the antenna design. It is relatively easy to develop a light weight S-band antenna. Indeed, current S-band patch antennas weigh only 80 g and it is foreseen than a 20 g patch antenna can be developed by simply removing the plate on the base of the antenna. However, there is not any light weight X-band antenna and its foreseen achievable weight would not be less than 100 g. In addition, the X-band antenna interface would most likely be a wave guide (heavier) whereas the S-band antenna interface would be an SMA cable (lighter).

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Finally, it must be pointed out that omni-directionality is very difficult to implement in X-band even impossible. If that was the case, the sub-system would need to be redesigned in order to provide omni-directionality for the uplink. The solution would likely be to have an X-band subsystem for payload telemetry (with an X-band antenna and requiring pointing the satellite during transmission) and one S-band sub-system for housekeeping telemetry and telecommands (with two S-band antennas providing an omni-directional pattern). Logically, that would lead into an increase of the mass and complexity of the whole communications sub-system. Taking into account the previous trade-off, the S-band is retained for NanoSat since it requires much simpler developments. However, the X-band must not be ruled out, especially if high data rate telemetry links were envisaged. 13.7.2 Mobile Phone Based Transponder For the NanoSat communications sub-system the use of a mobile phone based transponder (MPT) has also been considered. Since the first digital mobile phone appeared in 1991, mobile phones have spread all over the world. Nowadays mobile technology has largely evolved thanks to the push of commercial interests. One of the key points has been miniaturisation and power efficiency. The use of mobile phone technology for satellite communications has been assessed in several studies in ESA. So far this technology has been successfully tested on LEO satellites using both OrbCom and Iridium networks although at low data rate. Figure 13-8 shows what a mobile phone based communications sub-system would look like:

Mobile phone based transponder DHU HDR Telemetry transmitter + HPA

Figure 13-8: Mobile phone based transponder

In this configuration the DHU communicate in a typical way with the communication subsystem, i.e. sending and receiving digital information. Housekeeping telemetry and telecommands would be sent and received by the MPT while a dedicated transmitter, including an HPA, would deal with the payload telemetry. The reason of this division is that an MPT would only be able to transmit data up to 300kbps, thus for high data rate telemetry an additional transmitter is required. Thanks to this configuration, the mission would take advantage of the existing terrestrial mobile phone network for sending and receiving HKTM and TC, thus reducing the operational costs. It would also provide NanoSat with secured communications and the use of spread spectrum would prevent inter satellite interferences. Finally, mobile phone technology offers a solution that has already been optimized in mass and power and their chips are COTS due to the big number of

NANOSAT CDF Study Report: CDF-84(A) February 2009 page 195 of 267 mobile units produced every day in the world. The MPT would add about 100g to the mass budget. However, this increase of mass would be compensated by the suppression of the DW- converter in the nominal configuration. However, this solution has some drawbacks, which makes it less suitable than the nominal one: • Hardware: A mobile phone cannot be used as it is, a hardware development would be required, especially to adapt this technology to the space environment. • Frequency: The frequency allocated by the ITU for mobile phone communications is not applicable for space-to-Earth transmissions. In order to use the existing terrestrial mobile phone network the frequency regulations would need to be modified. • Antenna pattern: The antennas used in mobile base stations are conceived to radiate in the horizontal plane. The gain in the vertical direction can go down to 20dB. This would require the use of a high gain antenna in the S/C. A high gain antenna would produce a power flux density on the Earth surface above the allowed masks and. In addition, this kind of antenna would require a pointing mechanism, thus increasing the satellite mass and operations complexity. Finally, the footprint produced by the antenna would most likely cover several co-frequency mobile phone cells, producing interference with terrestrial services. The drawbacks presented by this solution make it less suitable than the solution that has been finally kept as the baseline. 13.7.3 Multifunctional Distributed Antennas An attractive possibility for Nano-satellites is the use of a number of RF exciters, distributed in suitable places on the external spacecraft surface. These exciters can be very small non-intrusive, for instance locally vertical monopoles (about 40mm at S-band) or slots (about the same size, but on the surface). They can, for instance, be located on the body wedges or between solar panels. The exciters act as feeder for the whole spacecraft structure, which effectively becomes the TT&C antenna. On-going TRP studies are showing that this concept is very well suited for spacecrafts of the size addressed in the NanoSat study when operating at S band. At X band the satellite becomes electrically too large to benefit from the distributed feeding concept. The drawback is the need of several pieces of RF cable to feed them all. However, the NanoSat study has show that the problem occurs only for the data download function, where higher link gain is required to achieve the desired data rate. The receive configuration remains therefore the same as found in traditional configurations, while the transmit configuration requires additional exciters. However, when looking at the RF power required it is clear that the SSPA will involve the combination of the output of several transistors. An interesting approach is then to remove the SSPA output combiner and feed each exciter from an individual transistor. Moving the power amplification stage close to the exciter would also allow to directly match the antenna to the transistor, further improving efficiency. Such approach significantly reduces the losses and simplifies the overall design.

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13.7.4 Electronics For Distributed Antennas Several further steps along this line of thought can be made. The first step consists in moving all the RF components, from the mixer upward, at the exciter locations, thus avoiding RF cables. The most advanced solution consists in using Direct Digital Synthesizers and AD Converters working directly at S-band (available commercially), very much along the same line of the mobile phone based transponder. Given the low data rate of the telecommands and the relatively high power consumption of fast ADCs it is likely better to first down-convert the Rx signal so as to use lower-speed ADCs. The exciters are the same irrespective of the solution adopted for the electronics. Clearly the fully digital option is very attractive for its simplicity and flexibility. The DHU would directly drive the distributed TT&C modules, composed of electronics and exciter. The electronics would consist of two hybrid chips (DDS and ADC) a small power conditioning unit and a command/telemetry circuit. The required filtering function can be easily integrated in the exciter, which has an intrinsic narrow-band characteristic given its small electric size.

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14 THERMAL 14.1 Requirements and Design Drivers The main design drivers to be used on NanoSat are: • Lightness • Simplicity • Reliability • Modularity. 14.2 Sub-System Description The following modules have been identified for the NanoSat thermal subsystem: • Black paint • MiSER – Miniature Satellite Energy Regulating Radiator • Thin Plate Heat Switch • Heater line (Kapton heaters + temperature sensors) • Heat pipe • MLI blanket • Heater with embedded temperature sensor. Some of the modules are only listed for completeness and to prove their existence, but actually not used in this assessment. 14.2.1 Interfaces MiSER and Thin Plate Heat Switch work coupled. Heater lines and Heaters with embedded temperature sensors interface with Power and DHS subsystems. 14.3 Module Descriptions 14.3.1 Module 1 (Black Paint) 14.3.1.1 Description and Overview Black paints are characterized by having a ratio alpha/epsilon ≅ 1. They are used especially to radiate the heat dissipated and homogenize the temperature of electronics boxes. 14.3.1.2 Key characteristics It has a density of 0.09 kg per unit area (m2). 14.3.1.3 Scalability Black paint is used on the surfaces used as radiator and on internal units for temperature homogenization.

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14.3.1.4 Multi mission suitability Black paint is used as surface finishing in basically all the spacecrafts. 14.3.1.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware. 14.3.2 Module 2 (MiSER – Miniature Satellite Energy Regulating Radiator)

Figure 14-1: MiSER 14.3.2.1 Description and Overview The MiSER Radiator has been developed for small spacecraft thermal control. The design consists of a flat radiator panel integrally mounted on a heat switch. The MiSER is mounted to the exterior of the spacecraft with the integral heat switch coupled to the heat load. The heat switch provides passive thermal control to the internal components of the spacecraft. The design is modular with each panel capable of dissipating 10 Watts. Single panels can be used to control the temperature of specific components or multiple panels can be used to control the temperature of the entire spacecraft. Custom radiator shapes provide the ability to match unusual geometries. When the temperature of the spacecraft rises above the set point temperature, the switch conductance increases, allowing the excess heat to be transferred through the switch to the radiator and out to space. When the temperature of the spacecraft drops below the set point temperature, the switch conductance decreases, insulating the spacecraft from the colder radiator panel and allowing the S/C devices to stay warm using a low level of device standby power. 14.3.2.2 Key Characteristics Technical specifications are reported below.

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Table 14-1: MiSER technical specifications 14.3.2.3 Scalability Different MiSER can be installed on the spacecraft according to the power that need to be evacuated. 14.3.2.4 Multi mission suitability This hardware can be used for any kind of mission that requires a heat switch due to power constraints. 14.3.2.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware.

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14.3.3 Module 3 (Thin Plate Heat Switch)

Figure 14-2: Thin Plate Heat Switch 14.3.3.1 Description and Overview A thin plate heat switch is a device for passively controlling the temperature of electronics and instrumentation on satellites and spacecraft. These devices represent a new family of tools for spacecraft thermal control. They can potentially reduce spacecraft power requirements while providing improved control and better reliability at a lower cost than conventional thermal control schemes. The devices provide a variable conduction heat path from the warm electronics or instruments to which they are mounted, to a cold panel or cold sink. The temperature of the electronics is controlled by the passive change in thermal conductance of the heat switch. The thermal conductance is adjusted internally and passively based on the temperature of the warm side of the switch. This self-regulating design allows for precise, reliable temperature control. In a typical application, the thin plate would be mounted between an instrument (or electronics box) and a cold sink such as a panel that radiates to space. When the electronics get too warm, the switch conductance increases, allowing the excess heat to be transferred through the switch to the cold sink. When the electronics get too cold, the switch conductance decreases insulating the electronics, allowing them to stay warm using their own heat or a low level of standby power. 14.3.3.2 Key Characteristics Mass is 20 grams. The other specifications are reported below.

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Table 14-2: Thin plate heat switch technical specifications 14.3.3.3 Scalability The number of heat switches to be used depends on how many MiSER radiators are needed on the satellite. 14.3.3.4 Multi mission suitability This hardware can be used for any kind of mission that requires a heat switch due to power constraints. 14.3.3.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware. 14.3.4 Module 4 (Heater Line – 2 Kapton Heaters + 1 Sensor) Kapton heaters and temperature sensors together create a heating line. 14.3.4.1 Description and Overview Kapton heaters: Kapton is a thin, semitransparent material with excellent dielectric strength. Kapton heaters (made in very small sizes) are ideal for applications with space and weight limitations, or where the heater will be exposed to vacuum, oil, or chemicals. Kapton heaters are radiation resistant to 106 rads if built with polyimide-insulated leads. Temperature range is -200/+200 C.

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Figure 14-3: Kapton heater

Temperature sensors: Different temperature sensors are available.

Table 14-3: Classification of temperature sensors Temperature sensor working principle is based on the electrical resistance variation that sensors have with temperature. With a constant value of current/voltage as input they give in output a voltage/current as function of the electrical resistance they have at that temperature. Repeatability/stability: Repeatability tells how well the sensor repeats subsequent readings at the same temperature. Stability is the absence of long term drift. For what concerns sensitivity, RTDs are preferred to thermistors, even if their sensitivity is the best. Thermistors have, in fact, a large negative TCR**. Resistance drops dramatically and non- linearly with temperature. They have a limited useful temperature range.

**Temperature Coefficient of Resistance [(R100 °C-R0 °C)/R0 °Cx100°C] RTDs represent the fastest growing segment among industrial temperature sensors and are largely used in space applications.

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An RTD sensing element consists of a wire coil or deposited film of pure metal. The element’s resistance increases with temperature in a known and repeatable manner. RTDs exhibit excellent accuracy over a wide temperature range. Contrarily to many sensors that have a high time lag, RTDs have a fast response and a time constant below 2 sec (time constant is defined as the time a sensor needs to reflect 63% of a step temperature change).

Figure 14-4: RTDs’ response vs. time 14.3.4.2 Key Characteristics Kapton heaters: Mass is about 35 grams; power that can be provided changes according to dimensions and mounting mean.

Figure 14-5: Kapton heater watt density vs. dimensions and mounting mean

RTDs as temperature sensors: RTDs elements for flat mounting (Discoil Thermal-Ribbons):

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Table 14-4: RTDs elements for flat mounting Typical weight is 2 to 5 grams. One Discoil Thermal-Ribbon measures and averages temperature readings over a surface of 0.19 m2. RTDs elements for pipe mounting (Strip Sensing Thermal-Ribbons):

Table 14-5: RTDs elements for pipe mounting Typical weight is 5 to 10 grams. Strip Sensing Thermal-Ribbons average temperatures along their length to eliminate point measurement errors. They can be also mounted on flat surfaces. 14.3.4.3 Scalability More than one heating line can be installed according to the needs (area to be warmed up, total heating power required). 14.3.4.4 Multi mission suitability This module can be used for any kind of mission. 14.3.4.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware.

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14.3.5 Module 5 (Heat Pipe)

Figure 14-6: Scheme of a heat pipe 14.3.5.1 Description and Overview A heat pipe is a heat transfer mechanism that can transport large quantities of heat with a very small difference in temperature between the hotter and colder interfaces. Inside a heat pipe, at the hot interface a fluid turns to vapour and the gas naturally flows and condenses on the cold interface. The liquid falls or is moved by capillary action back to the hot interface to evaporate again and repeat the cycle. Some example fluids are water, ethanol, acetone, sodium, or mercury, chosen to match the heat pipe operating temperature. Inside the pipe's walls, a wick structure exerts a capillary pressure on the liquid phase of the working fluid. This is typically a sintered metal powder or a series of grooves parallel to the pipe axis, but it may be any material capable of exerting capillary pressure on the condensed liquid to wick it back to the heated end. Active control of heat flux can be obtained by adding a variable volume liquid reservoir to the evaporator section. Variable conductance heat pipes employ a large reservoir of inert immiscible gas attached to the condensing section. Varying the gas reservoir pressure changes the volume of gas charged to the condenser which in turn limits the area available for vapour condensation. Thus a wider range of heat fluxes and temperature gradients can be accommodated with a single design. 14.3.5.2 Key Characteristics In general heat pipes can be up to 3.5 meters long depending on the application. Mass is calculated as 0.35 kg/m. Power may need to be applied to keep the fluid far from its freezing point. Features of a typical heat pipe used on nanosat are hereafter reported.

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Table 14-6: Heat pipe characteristics 14.3.5.3 Scalability More than one heat pipe can be applied depending on the total heat that needs to be evacuated and the area this heat needs to be spread on. 14.3.5.4 Multi mission suitability This module can be used for any kind of mission. The materials chosen as working fluid depend on the temperature conditions in which the heat pipe must operate, with coolants ranging from liquid helium for extremely low temperature applications (2–4 K) to mercury (523–923 K) & sodium (873–1473 K) and even indium (2000–3000 K) for extremely high temperatures. The vast majority of heat pipes for low temperature applications use some combination of ammonia (213–373 K), alcohol (methanol (283–403 K) or ethanol (273–403 K)) or water (303–473 K). 14.3.5.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware. 14.3.6 Module 6 (MLI Blanket)

Figure 14-7: MLI blanket

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14.3.6.1 Description and Overview Multi-layer insulation or MLI is mainly intended to reduce losses by thermal radiation. The principle behind MLI is radiation balance. To see why it works, start with a concrete example - imagine a square meter of a surface in outer space, at 300 K, with an emissivity of 1, facing away from the Sun or other heat sources. From the Stefan-Boltzmann law, this surface will radiate away 460 watts of power. Now imagine we place a thin (but opaque) layer 1 cm away from the plate, thermally insulated from it, and also with an emissivity of 1. This new layer will cool until it is radiating 230 watts from each side, at which point everything is in balance. The new layer receives 460 watts from the original plate. 230 watts is radiated back to the original plate, and 230 watts to space. The original surface still radiates 460 watts, but gets 230 back from the new layers, for a net loss of 230 watts. So overall, the radiation losses have been reduced by half by adding the additional layer. More layers can be added to reduce the loss further. The layers of MLI are usually made of very thin plastic, about 10 micrometres (1/4 mil) thick, such as Mylar or Kapton, coated on one side with a thin layer of metal on both sides, typically silver or aluminium. 14.3.6.2 Key Characteristics MLI weighs 0.5 kg/m2 considering margin to allocate some mass to struts, Velcro. No power is allocated for this unit. Dimensions as well as shape can be adjusted according to the needs. 14.3.6.3 Scalability Performances can be improved by augmenting the number of layers. Ideally, a N layer blanket reduces the thermal radiation to 1/(N+1) of the original value. However above a certain limit (30-35 layers) it is not worth increasing this number as performances does not increase. 14.3.6.4 Multi mission suitability This module can be used for any kind of mission. 14.3.6.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware. Only adaptation case by case is needed (number of layers, finishing, shape).

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14.3.7 Module 7 (Heater With Embedded Temperature Sensor)

Figure 14-8: Heater/sensor 14.3.7.1 Description and Overview Integrated heater/sensors are the ideal solution for many temperature control problems. Combining an etched foil heating element with an accurate, stable RTD or thermistor sensor in a single package provides a reliable system with reduced parts count and simplified installation. Standard heater/sensors have the sensor element located in a non-heating area to measure the heat sink temperature — not the heating element temperature. The result is a more accurate reading and better control. Precise location of the sensor ensures consistent readings every time. 14.3.7.2 Key Characteristics Mass is about 50 grams; power that can be provided changes according to dimensions and mounting means.

Figure 14-9: Heater/sensor watt density vs. dimensions and mounting means

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14.3.7.3 Scalability More than one heater/sensor unit can be installed according to the needs (area to be warmed up, total heating power required). 14.3.7.4 Multi mission suitability This module can be used for any kind of mission. 14.3.7.5 Development Plan No development plan is needed as the unit is already an off-the-shelf thermal hardware. 14.4 Scenario Study Case 1 (LEO Scenario) Three orbits have been considered. Orbit 1: h = 600 km RAAN = 0.00 deg i = 97.06 deg S/C x-axis Nadir-pointing S/C z-axis Normal to orbit

Figure 14-10: LEO Scenario – Orbit 1 Orbit 2: h = 600 km RAAN = 90.00 deg i = 97.06 deg S/C x-axis Nadir-pointing S/C z-axis Normal to orbit

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Figure 14-11: LEO Scenario – Orbit 2 Orbit 3: h = 600 km RAAN = -30.00 deg i = 97.06 deg S/C x-axis Nadir-pointing S/C z-axis Normal to orbit

Figure 14-12: LEO Scenario – Orbit 3 14.4.1 Design Drivers The thermal subsystem shall guarantee that in every phase of the mission the units mounted on the spacecraft stay within their temperature ranges (say -25 / + 50 degC).

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14.4.2 Module Selection (Justification) The satellite has body mounted solar cells on the majority of all its surfaces. This helps produce the required power in all the possible attitudes. The rest of the spacecraft is black painted for temperature homogenization (internally) and for radiative purposes (externally). Battery is wrapped in MLI tents and kept warm in eclipse by using heating lines. The results of a preliminary and very simplified model show that using only black paint the satellite can cope with the requirements. Orbit 1:

Instrument on 6 AOCS on 2 DHS on 3 Comms on 4 Figure 14-13: Temperature evolution during one orbit for Orbit 1 Orbit 2:

Instrument on 6 AOCS on 2 DHS on 3 Comms on 4 Figure 14-14: Temperature evolution during one orbit for Orbit 2

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Orbit 3:

Instrument on 1 AOCS on 2 DHS on 3 Comms on 4 Figure 14-15: Temperature evolution during one orbit for Orbit 3

14.4.3 Mass/Power Budget

Table 14-7: Mass budget LEO case 2 W are allocated for batteries temperature regulation to avoid that when in eclipse its temperature falls below zero. 14.5 Scenario Study Case 2 (LEO with Propulsion) For the thermal analysis, this case is identical to Study Case 1. 14.6 Scenario Study Case 3 (GTO Scenario) One orbit has been considered: perigee = 250 km apogee = 35143 km inclination = 6 deg S/C x-axis Sun-pointing

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S/C z-axis Normal to orbit

Figure 14-16: GTO Scenario 14.6.1 Design Drivers The thermal subsystem shall guarantee that in every phase of the mission the units mounted on the spacecraft stay within their temperature ranges (say -25 / + 50 C). 14.6.2 Module Selection (Justification) Due to the long eclipse the satellite has to be shielded from external environment by using MLI blankets. Three panels are left free to radiate to space (panels 5,9 and 10). On the body mounted part of the solar array MLI are used on its rear side to decouple it from the S/C structure. Because the units are mounted on different panels (Instrument on 1, AOCS on 2, DHS on 4, Comms on 5, HPs have to be used as bridge between unit mounting plates and radiator. Black paint is used internally for homogenization purpose. Battery is wrapped in MLI tents and kept warm in eclipse by using heating lines. A preliminary and very simplified model shows that this design can cope with the requirements (with additional measures as below).

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150 Top panel Big lateral panel 100 BM-SA Big lateral panel 50 Big lateral panel Bottom panel Small lateral panel 0 Small lateral panel 0 5000 10000 15000 20000 25000 30000 35000 40000 Small lateral panel -50 Small lateral panel SA1 front Temperature [C] SA2 front -100 MLI out top panel MLI out big lat panel -150 MLI on BM-SA rear side MLI out big lat panel MLI out bottom panel -200 Time [sec]

Figure 14-17: Temperature evolution during one orbit Based on the results of the thermal analysis, it is however noted that further measures are required in order to satisfy the Star Tracker temperature requirements. Those proposed are: • To place the Star Tracker on one of the small lateral panels facing away from the Sun (i.e. panel 9 or 10) • To thermally decouple the optical head.

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14.6.3 Mass/Power Budget

Table 14-8: Mass budget GTO case 5 W are allocated for batteries temperature regulation to avoid that when in eclipse its temperature falls below zero.

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15 GROUND SEGMENT & OPERATIONS 15.1 Requirements and Design Drivers The requirements for the NanoSat ground segment were given as: • 3.4 m ground antenna • 2 Mbps dump rate • Low cost operations concept • Support for GTO and LEO missions. 15.2 Assumptions and Trade-Offs This CDF is a special case because the aim is not to find the most optimum ground segment for a particular mission but to create an end-to-end system that fits as many missions as possible. Therefore the first step is to challenge the requirements regarding the size of the ground antenna and dump rate. These requirements were reformulated as follows “Identify a low cost operations concept that is applicable to the largest range of NanoSat mission possible”. 15.2.1 Ground Station Selection Trade Three different ground station positions were considered in the trade: Polar, Mid-European and Equatorial. Polar stations have good coverage for LEO missions, Equatorial stations have good coverage for GTO missions but neither can be used for both types of mission. Mid-European stations e.g. Redu, can be used for either orbit but are not optimum. For a LEO mission Redu sees the satellite 5-6 times a day and has gaps of up to 11 hours. For GTO missions, Redu sees that satellite every day but only at large range distances (more than 15,000 km). 15.2.2 Network Trade If a network (e.g. ESTRACK) is used then the problem with ground station selection does not exist as the best ground station for the orbit type can be chosen. However network stations are rented by the hour therefore this can be expensive for certain types of mission. In trying to satisfy the requirement for low cost operations the possibility of buying and operating a small (3.4 m) antenna was investigated. This is the PROBA-1 approach. The most important thing to understand regarding the PROBA-1 approach is that the initial investment in the antenna must be balanced by the lower operations and maintenance costs over the mission. This means it only works if the mission is long enough and that it is co-located with a control facility e.g. Redu or ESOC. The use of a small antenna and mid European latitudes will restrict the data return. 15.2.3 Concept Trade The conclusion is that no single ground station or concept is perfect for all possible mission requirements. Therefore a menu of possible concepts is proposed from which the best can be selected depending on the data return and timeliness requirements. • Small ‘a la PROBA’: Small antenna, co-located with control facility in Europe • Scalable: TT&C using small antenna in Europe. Data return using infrequent dumps at 15 m antennas of a professional network.

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• Classic: TT&C and data return using a professional network.

Timeliness

Classic 5 passes/day 11 hour gaps

Small Scalable Data return XYZ

Figure 15-1: Concept v Mission Requirements Only the classic mission concept can be chosen if the timeliness requirements do not allow 11 hours gaps (LEO). However it also shows that if the timeliness requirements are relaxed then there are a range of mission concepts that are dependant on data return. “X’ corresponds to the data volume that can be returned using a small antenna for 5 passes a day, “Y” to the data volume that can be returned using a large antenna twice a day and “Z” to the data volume that can be returned using a large antenna 14 times a day (every orbit). The values of these parameters depend on the frequency and type of on-board antenna chosen. 15.2.4 Frequency Trade Two possible frequencies were investigated for NanoSat: X band and S band. S band has the advantage that the same frequency bands are applicable for two main probable applications; Earth Observation and Space Science. For X band these are different which means extra modules would have to be provided on the spacecraft for the different applications. However X band has an enormous advantage in terms of bandwidth available. The use of S band is restricted to 6 MHz per spacecraft. X band is limited to 10 MHz for space science and 375 MHz for earth observation. This extra bandwidth can be exploited to produce much higher data returns for the same concepts. The limited bandwidth available in S band can be optimised by using better filters and variable coding and modulation schemes. However this can also be applied to X band and does not fundamentally change the advantage.

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S band X band Redu 3.4 m Gbits Gbits Average (no removal of bad passes) 1.1 1.1 Kiruna 15 m Average (no removal of bad passes) 13 61 Good pass (2-3 a day) 16 82 Best pass 16.5 91.5

Table 15-1: Frequency v Data return The table above was calculated with the following assumptions: • LEO SSO orbit at 600 km • 2W transmitter power on-board • Low gain antenna on-board (-3dB assumed in all cases) • Variable coding and modulation (VCM), changing with elevation to ground station • Dumping starts and ends at 5 deg elevation. The table shows that one good pass using the 15m antenna at Kiruna (available two or three times per day) in S band we can dump the equivalent of 15 times the data of one pass with a 3.4m antenna. This is impressive until you see that equivalent in X band is 75 times.

Data return v antenna size/frequency

s 100 80 60 40 20

Data Return Gbit Return Data 0 Once 4 times S-band X-band 3.4m 3.4m 15m 15m Passes

Figure 15-2: Data return v antenna size and frequency When you take into account that such passes only cost a low hour rate then the economics become clear. The cheapest way to return data from NanoSat is using infrequent dumps, large antennas and X band. The frequency trade can also be visualised in terms of the slopes of the cost/bit returned curve. Obviously the more data to be returned, the greater the advantage of using X band.

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Cost v Data Return

) 10 8 S band X band 6 4

Cost (KEuros 2 0 0 500 1000 1500 Data Return (Gbits)

Figure 15-3: Frequency Trade in cost terms for 15 m antenna 15.2.5 Scalability Trade Now we have performed the frequency trade we can enter the parameter values in Figure 15-1.

1700 wit h isoflux? X band Much more with rotation? 160 850

S band

4.5 32 180 Data return (Gbits/day) ‘Small’ ‘S calable’ ‘Classic’ Li mi t Limit Limit 4 pass/day 2 pass/day 14 pass/day

Figure 15-4: Scalability Trade for LEO This shows that a real advantage of X band is that the low cost scalable operations concept can be applied for a much wider range of missions than for S band. In fact the scalable concept is still valid for X band almost up to the limit of the classic mission in S band (160 compared to 180 Gbits/day). Also the X band option provides NanoSat with a much higher capability in terms of data return. To put things in perspective these figures are similar to the Cryosat-2 dump capability.

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Mission Data Return (Gbits/Day) GOCE 0.7 AEOLUS 5.2 SMOS 15 CYROSAT-2 400 EARTHCARE 600 Table 15-2: ESA EO missions v Data return Finally, it should be noted that while S band effectively hits a hard limit at 180 Gbits/day, the X band limit of 850 Gbits/day is not hard. It can be further leveraged by exploiting extra power, combined with VCM techniques. Where will this power come from? The most effective improvement could be the use of an X band isoflux antenna if the mission was nadir pointing. Calculations indicate this could be used to double data return to 1700 Gbits/day for the classic concept and raise the limit of the scalable concept to 320 Gbits/day. Alternatively the spacecraft could be actively pointed during the pass (similar to PROBA-1 when it takes a picture). 15.2.6 Ground Station Equipment Trade The baseline small mission concept means buying and operating a 3.4 m tracking S band antenna. This can be bought for a moderate cost. The scalable concept could be cheaper. This is because it is not being used to dump payload data but only housekeeping telemetry. A 3.4 m antenna would not be required to handle the very low data rates required. This could be exploited to make cost savings in one of three ways: • Use a small antennas produced for radio amateurs (low cost) • Use spread spectrum techniques and ground station antennas with no moving parts (or arrays of these antennas) • Use space based mobile phone access (there is a GSP study on-going). Note this would be unbeatable in terms of providing access. 15.3 Baseline Design The baseline design is a ground control concept using a 3.4m, S band antenna, co-located with a operational control facility in Europe e.g. Redu. This system would meet the CDF requirements as stated. Timeliness requirements up to 11 hours gap and data return up to 4.5 Gbits/day could be met. Although we spend most effort investigating the ground segment from a ground station point of view it should be recognised that most operations costs are associated with the mission requirements themselves. If a mission has challenging data return requirements, complicated mission planning and tight reaction times then it will not be cheap to operate. To ensure the mission operational concept is as low cost as possible the following approach is assumed:

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• A risk-free safe mode is provided. This means it will be used in the mission. This allows the pre-launch testing of mission critical procedures/systems only • A “no LEOP” operations concept will be used. This means relying on safe mode, flexible schedules and industry experts for trouble shooting in the early stages • The industrial and operations teams will be closely integrated. This will reduce the knowledge transfer overhead. They will provide support for each other during peaks in workload • A low cost operations team and infrastructure is assumed to already exist. NanoSat will then be absorbed into it • No mission planning or time critical events will be part of the mission • Only best effort requirements for data return, reaction, turnaround times and ground system availability will be required. Nanosat control will be integrated into an existing infrastructure of a least four people. The delta required in terms of manpower will be one man per year. Operations preparation is expected to start at launch minus two years. Operations will be automated as far as possible. No manning will be provided at night, weekends or holidays. 15.4 Options Although the baseline is the small mission approach, NanoSat could be deployed as low cost mission with very respectable mission data returns by using VCM, X band and isoflux antennas. The scalable mission operations concept would also allow other novel operations concepts to be employed: • Pre-processing: Take lots of pictures, and downlink ‘quick check’ information (e.g. thumbnails) through the TT&C link. Delete those not required by telecommand. This can reduce the downlink cost even further • Opportunity: Only take data when a special opportunity arises but then take as much as you can. Alert the ground that a dump is required through the TT&C link. In this way you only pay for useful data • Post processing: Reduce planning complexity by taking much more data than you actually need. 15.5 Technology Requirements The following technology is required or would be beneficial to this domain: • Mass market baseband equipment supporting VCM at low bit rates. A high bit rate CORETX is expected on the market in 2 years so this is regarded as a safe assumption.

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16 DEVELOPMENT RISK 16.1 The Risk Assessment Process Risk management is an organized, systematic decision making process that efficiently identifies, analyzes, plans, tracks, controls, communicates, and documents risk in order to increase the likelihood of achieving the project goals. The procedure comprises four fundamental steps RD[26]: • Step 1: Definition of the risk management policy which includes the mission success criteria and the severity & likelihood categorizations • Step 2: Identification and assessment of risks in terms of likelihood and severity • Step 3: Decision and action (Risk acceptance or implementation of mitigating actions) • Step 4: Communication and documentation.

Figure 16-1: ECSS-M-ST-80C, 2008 Risk Management Process 16.2 Risk Management Policy The CDF risk management policy for Nanosat aims at handling risks which may cause serious cost, schedule, and/or technical impact on the project’s required module developments. The following actions have been carried out as a basis for the implementation of the risk management process: • Identification of the module development success criteria • Establishment of a scoring scheme for the severity of consequences of undesired events affecting the module development success criteria (1-3) • Definition of likelihood or probability of occurrence levels (A-C) • Establishment of a risk index scheme to indicate the magnitudes of the risks of the various risks scenarios • Establishment of criteria to determine the mitigating actions to be taken on risks • Establishment of a method for the ranking and comparison of the risks.

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16.3 Dependability and Safety Strategy The following dependability and safety strategy is proposed based on D/TEC’s PA & Safety Strategy for In-orbit Demonstration Lightsat Concept approach RD[27]: • Failure Tolerance o The capability to sustain single or multiple failures may be reduced or eliminated in accordance with Nanosat’s Risk Policy which accepts higher risk levels o Redundancy is not required or may not necessarily be required unless to prevent HW failures leading to catastrophic or critical consequences as defined in ECSS-Q-ST-30 o Redundancy can be provided at mission level with several Nanosat’s working in parallel • Limited failure propagation requirements (see PA & Safety Strategy for IOD Lightsat) • Reliability o High reliability components and units shall be used in order to achieve high S/C reliability without a redundant architecture. Short mission lifetime is a plus • Reliable Safe Mode • Component Quality levels may be traded against environmental factors (radiation) and mission duration • Dependability Analyses o Functional analysis only at system level o FMECA at I/F level to guarantee that there is no failure propagation between units o WCA, Parts De-rating Analysis only for critical components • Critical Items List (CIL) o Items with single point failures and severity classified as catastrophic or critical shall be listed • Dependability Testing o Verification of functioning and performance of the onboard FDIR. If ground intervention is required to perform FDIR for the flight segment, function and performance of these capabilities shall be verified end to end • Safety o Apply ECCS-Q-ST-40 (Tailoring TBD, no special reduction foreseen for the case of a Nanosat) • Safety goal: Identify, reduce to acceptable levels or eliminate all possible safety hazards during all mission phases. 16.4 Module Development Success Criteria Module development for Nanosat is considered successful when the technical, programmatic/schedule, and cost objectives are met. These are collected in Table 16-1:

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Domain Success Criteria Technical Module performance meets the requirements specification Schedule Module development at completion is within schedule Cost Module development cost at completion is within budget Table 16-1: Nanosat Mission Success Criteria 16.5 Module Development Severity Categorization The risk scenarios are classified according to their domains of impact. The consequential severity level of the risk scenarios is defined according to the worst case potential effect with respect to cost, schedule, and/or technical performance. Identified risks that may jeopardize and/or compromise Nanosat’s module development will be ranked in terms of likelihood of occurrence and severity of consequence. The scoring scheme with respect to the severity of consequence on a scale of 1 to 3 is established in Table 16-2, and the likelihood of occurrence is normalized on a scale of A to C in Table 16-3.

Score Severity Cost Schedule/Programmatic Technical 3 High Cost increase result in module Delay implies module Loss of module development cancellation. development cancellation. function.

2 Medium Cost requiring funding beyond TBD months needed to restore Major degradation plans within approved the schedule. of module margins. performance.

1 Low Cost increase requiring some TBD weeks needed to restore Minor degradation redeployment of resources. the schedule. of module performance

Table 16-2: Severity Categorization

Score Likelihood Definition

C High The event is certain to occur, will occur once or more times per project.

Will occur frequently, about 1 in 10 projects. B Medium Pf=0.1 R=0.9

Will occur seldom, about 1 in 1000 projects A Low Pf=0.001 R= 0.999 Table 16-3: Likelihood Categorization

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16.6 Risk Index A Risk Index is given as a combination of the likelihood of occurrence and the severity of consequence for a given risk item. The risk Index chart is represented in Figure 16-2. Severity 3 3A 3B 3C 2 2A 2B 2C 1 1A 1B 1C A B C Likelihood Figure 16-2: Risk Index Chart For Nanosat, risk ratings of low (green), medium (yellow), and high (red) will be assigned based on the criteria of the Risk Index Scheme. The level of criticality for a risk item is denoted by the analysis of the risk index. Following the scheme described in the previous section the highest possible Risk Index will therefore be 3C, and the lowest possible Index, 1A. 16.7 Top Risk Log The risk log is organized by subsystem and summarized in the Top Risk Index Charts below:

AOCS 9 MODULES POWER 3 MODULES

Severity Severity 3 15 5 3 03 0 2 09 1 2 13 0 1 0 2 0 1 0 0 0 A B C A B C Likelihood Likelihood

COMMS/ANTENNA 7 MODULES PROPULSION 7 MODULES

Severity Severity 3 20 2 3 03 0 2 09 2 2 13 4 1 05 0 1 01 0 A B C A B C Likelihood Likelihood

DHS 4 MODULES STRUCTURES 6 MODULES

Severity Severity 3 02 2 3 01 1 2 04 0 2 10 0 1 00 0 1 30 0 A B C A B C Likelihood Likelihood

MECHANISMS 4 MODULES THERMAL 7 MODULES

Severity Severity 3 03 0 3 01 0 2 12 0 2 03 0 1 00 0 1 30 0 A B C A B C Likelihood Likelihood Table 16-4: Top Risk Index Charts by Subsystem

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 AOCS Digital Sun 2B Increase in estimated Cost/Schedule Expensive EQM with includes Prototype TRP is currently Apply contingency margins Expected to reach TRL 6 Sensor cost, lack of funding, or /Programmatics component qualification. funded and running. in development schedule. by 2010. Expected to reach possible delays in Complex qualification Apply margins in cost TRL 9 by 2013. schedule. procedure. budget to avoid overrun.

Digital Sun 2B Degradation of module Technical Radiation sensitivity. Radiation issues are to be Invest in technology and Verify compliance with 40 Sensor performance. resolved during testing. Krad tolerance qualification testing. requirement. Digital Sun 3B Critical technology Technical Power conditioning, pixel Prototype testing should If development activities Invest in technology and Sensor issues affecting design and manufacturability resolve technical risks. are successful risk should testing. technical performance. are current technical concerns. be moderated.

Star Tracker 3B Critical technology Technical Critical areas in optics Pre-development activities If development activities Invest in technology and issues affecting technology and chip (2009-2012) geared are successful risk should testing. technical performance. integration. towars resolving technical be moderated. risks.

Star Tracker 2B Increase in estimated Cost/Schedule IP core purchase for EM. Apply contingency margins Expected to reach TRL 6 Expected to reach TRL 9 cost or possible delays /Programmatics in schedule. by 2013. by 2016. in schedule. Star Tracker 1B Increase in estimated Cost Expensive EQM with includes Cost risk is regarded by Apply margins in cost cost or lack of funding. component qualification. specialist as minor (low budget to avoid overrun. severity). Star Tracker 3C Possible delays in Cost/Schedule TRL 2 (2009). Complex Implement contingency Expected to reach TRL 6 Expected to reach TRL 9 schedule. /Programmatics qualification procedure. margins in schedule. by 2013. by 2016. Gyro 3B Critical technology Technical Controller chip is a complex Adequate radiation The detector chip is Invest in technology and issues affecting mixed signal ASIC. There are hardening procedure. already at TRL 7. testing. technical performance. uncertainties regarding its Verify compliance with 40 perfomance in the space Krad tolerance environment. requirement. Earth Sensor 3C Critical technology Technical Uncertainties in the technical An activity is currently If activities are successful Invest in technology and issues affecting concept. underway to demonstrate risk should be moderated. testing. technical performance. the proof of concept and key characteristics.

Earth Sensor 3C Possible delays in Schedule Technology is at TRL 1 (2008) Invest in development and Expected to reach TRL 3 Expected to reach TRL 7 schedule. /Programmatics testing. by 2010. by 2012. GNSS 3C Critical technology Technical Uncertainties in the use of the Technology development Verify compliance with 40 Invest in technology and Receiver issues affecting existing ground technology for required to redesign Krad tolerance testing. technical performance. use space applications system for space requirement. (radiation issues). environment (rad hard)

GNSS 2B Increase in estimated Cost/Schedule Uncertainties in budget or time Need for comercial or Invest in technology and Receiver cost, lack of funding, or /Programmatics allocation for the required scientific applications of testing. possible delays in technology development. Nano space technology. schedule. Magnetometer 2B Critical technology Technical Major development required in Detector chip currently Miniturize electronics onto Invest in technology and issues affecting the electronics. exists. a rad hard mixed signal testing. technical performance. ASIC. Verify compliance with 40 Krad tolerance requirement. Magnetometer 3B Increase in estimated Cost/Schedule Uncertainties in budget or time Need for comercial or Invest in technology and cost, lack of funding, or /Programmatics allocation for the required scientific applications for testing. possible delays in technology development. this Nano space schedule. technology. Magnetorquer 2B Critical technology Technical Delopment required in the Development not expected Chip to be manufactured in Invest in technology and issues affecting production of a current control to be an issue. the prototype model and testing. technical performance. chip capable of handling qualified in the EQM. 0.25Amps at 5v. Magnetorquer 2C Increase in estimated Cost/Schedule Uncertainties in budget or time Need for comercial or Invest in technology and testing. cost, lack of funding, or /Programmatics allocation for the required scientific applications of possible delays in technology development. Nano space technology. schedule. Reaction 3A Critical technology Technical Motor & motor controller Main development Micro RW motor concepts Exact design approach to Wheel issues affecting technology is considered required during prototype are already existing. the motor to be decided via technical performance. critical. phase where the mixed experimentation in the signal rad hard ASIC predevelopment. motor controller will be developed.Verify compliance with 40 Krad tolerance requirement.

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 AOCS Reaction Wheel 2B Increase in estimated Cost/Schedule Uncertainties in budget or time Need for comercial or Invest in technology and testing. cost, lack of funding, or /Programmatics allocation for the required scientific applications of possible delays in technology development. Nano space technology. schedule. Navigation 3B Critical technology Technical Critical areas in optics Pre-development activities Invest in technology and Camera issues affecting technology and chip (2009-2012) geared testing. technical performance. integration. towars resolving technical risks.

Navigation 2B Increase in estimated Cost/Schedule IP core purchase for EM. Apply contingency margins Expected to reach TRL 6 Expected to reach TRL 9 Camera cost or possible delays /Programmatics in development schedule. by 2013. by 2016. in schedule. Navigation 1B Increase in estimated Cost Expensive EQM with includes Cost risk is regarded by Apply margins in cost Camera cost or lack of funding. component qualification. specialist as minor (low budget to avoid overrun. severity) Navigation 3C Possible delays in Cost/Schedule TRL 2 (2009). Complex Implement contingency Expected to reach TRL 6 Expected to reach TRL 9 Camera schedule. /Programmatics qualification procedure. margins in development by 2013. by 2016. schedule. AOCS General 2B Critical technology Technical EMC/EMI issues with Appropriate EMC/EMI Invest in technology and issues affecting miniturization. testing to avoid or testing. technical performance. minimize risk.

Table 16-5: Top Risk Log AOCS Modules

Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 COMMS/Antenna X/S band 3A Critical technology Technical Development required for Some prototypes of class Two TRP activities Invest in technology and UP/DW issues affecting Class F SSPA (70% F SSPAs have been currently ongoing at ESA testing. Converter technical performance. efficiency). Technical developed and for Class F SSPA. Efficient Power complexity of integrating up successfully tested in However, products not Amplifier conversion, down conversion, ground environment. applicable to NanoSat. low noise amplifier and SSPA in the same unit.

X/S band 3A Critical technology Technical Development required for GaN Several activities going on Invest in technology and testing. UP/DW issues affecting technology. Technical at ESA to implement Converter technical performance. complexity of combining Class SSPA using GaN Efficient Power F SSPA with GaN technology. technology instead of Amplifier Issues: Temperature GaAs. dissipation must be carefully assessed. Temperature constraints are more relaxed when using GaN.

X/S band 2B Increase in estimated Cost/Schedule Prototype currently at TRL 3. Expected to reach TRL 9 Prototype TRP is currently Apply contingency margins UP/DW cost , lack of funding, /Programmatics Four UP/DW converters have by 2013 funded and runnning. in development schedule. Converter or possible delays in to be developed each one Apply margins in cost Efficient Power schedule. providing a different RF output budget to avoid overrun. Amplifier power level level (1W, 2W, 4W Invest in technology and and 8W) testing. Mobile Phone 3C Critical technology Technical Important technology The advantages and Signal should be shifted to Invest in technology and Based issues affecting development required to disadvantages of using an allocated frequency in S- testing. Transponder technical performance. integrate mobile technology existing mobile phone band. with the necessary power technology has been amplification for a Space-to- assessed in several ESA Earth transmission. Space to activities. Mobile Earth transmission at high technology using data data rate has never been relay (like IRIDIUM) has tested. Mobile technology is been successfully tested. not space qualified plus mobile This means space-to- frequencies can not be used space transmission at low for a space-to-Earth data rate. transmission.

Mobile Phone 3C Increase in estimated Cost/Schedule TRL 2 in 2008. No Invest in technology and Apply contingency margins Apply margins in cost Based cost, lack of funding, or /Programmatics development activities are testing. in schedule. budget to avoid overrun. Transponder possible delays in being carried out at the schedule. moment for this technology. Lightweight S- 1B Increase in estimated Cost/Schedule The activity involves a This development poses No changes to technology Apply margins in cost Band Antenna cost, lack of funding, or /Programmatics modification of the mechanical no major technical risks. or other design elements budget to avoid overrun. possible delays in design and interfaces of an are required. Apply contingency margins schedule. existing flight antenna to in development schedule. reduce mass.Technology at Invest in technology and TRL 6 in 2008. Expected to testing. reach TRL 9 by 2011.

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 COMMS/Antenna Lightweight X- 1B Increase in estimated Cost/Schedule Currently at TRL 3 with a high Expected development This development poses no Apply margins in cost band Low Gain cost, lack of funding, or /Programmatics estimated cost of around 800K time of 4 years. major technical risks. No budget to avoid overrun. Antenna possible delays in Euros. changes to technology or Apply contingency margins schedule. other design elements are in development schedule. required. Invest in technology and testing. Multifunctional 2B Critical technology Technical/cost Reliability mass and cost Use of separate units with Apply margins in cost Invest in technology and Distributed issues affecting impact of RF and mechanical standard blindmate budget to avoid overrun. testing. Antenna technical performance. interface between antenna and connectors. System Increase in estimated electronic module. cost.

Multifunctional 1B Critical technology Technical Uncertain feasibility of Include separate filter. Integrate filter with Invest in technology and Distributed issues affecting integrating the filter functon electronics module. testing. Antenna technical performance. within the antenna. System Multifunctional 2B Critical technology Technical Mass impact and overall Fall-back option is not to Invest in technology and testing Distributed issues affecting complexity of integrate the antenna with Antenna technical performance. structrure/antenna interfaces the structure. System towards the platform.

Multifunctional 1B Critical technology Technical/Cost Mass and cost impact Use fully metallic solution Use of fully composite Mixed metallic/composite Distributed issues affecting depending on the technology incase for cost solution for mass solution could be a good Antenna technical performance. selected: Fully metallic (higher considerations. considerations. compromise for both mass System Increase in estimated mass/lower cost), composite and cost. cost. (lower mass/higher cost) or mixed. Multifunctional 2C Increase in estimated Cost/Schedule TRL 1 in 2008. Expected to Invest in technology and Apply margins in cost Apply contingency margins Distributed cost , lack of funding, /Programmatics reach TRL 9 by 2015. Major testing. budget to avoid overrun. in development schedule. Antenna or possible delays in technology development System schedule. required. Electronics for 2B Critical technology Technical Mass and complexity impact of Use down-conversion to Invest in technology and Distributed issues affecting very high gain Rx modules. avoid oscillations. testing. Antenna technical performance. System

Electronics for 2B Critical technology Technical Mass and size impact of the Accept larger units (may Invest in technology and Distributed issues affecting miniturization. geopardize its application testing. Antenna technical performance. to Nanosat) System

Electronics for 2B Critical technology Technical Technical feasibility of the Include separate filter. Invest in technology and Distributed issues affecting integration of the filter function testing. Antenna technical performance. (if excluded from the antenna). System

Electronics for 2B Critical technology Technical Mass impact and overall Separate module from Invest in technology and Distributed issues affecting complexity of interfaces with antenna. testing. Antenna technical performance. antenna. System

Electronics for 2C Increase in estimated Cost/Schedule TRL 1 in 2008. Expected to Invest in technology and Apply margins in cost Apply contingency margins Distributed cost , lack of funding, /Programmatics reach TRL 9 by 2015. Major testing. budget to avoid overrun. in development schedule. Antenna or possible delays in technology development System schedule. required.

Comms 2B Critical technology Technical EMC/EMI issues with Appropriate EMC/EMI Invest in technology and (General) issues affecting miniturization. testing to avoid or testing. technical performance. minimize risk.

Serially 2B Increase in estimated Cost/Schedule Low technology readiness A lab demonstrator already Standard is already an Invest in technology and Concatenated cost , lack of funding, /Programmatics level. exists and has been tested orange book in CCSDS testing. Convolutional or possible delays in successfully. and it is expected that it Code (SCCC) schedule. becomes green book in some years. Low Data Rate 1B Increase in estimated Cost/Schedule Development required. Invest in technology and Apply margins in cost Reciever for cost , lack of funding, /Programmatics testing. budget to avoid overrun. Ground Station or possible delays in Apply contingency margins schedule. in development schedule.

Table 16-6: Top Risk Log COMMS/ANTENNA Modules

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 DHS General 2B Critical technology Technical EMC/EMI issues with Appropriate EMC/EMI Invest in technology and Purpose issues affecting miniturization. testing to avoid or testing. Interface ASIC technical performance. minimize risk. (RT ASIC)

General 2B Increase in estimated Cost/Schedule High estimated non recurring Use of mature IPs. Low Apply margins in cost Purpose cost , lack of funding, /Programmatics cost. development complexity. budget to avoid overrun. Interface ASIC or possible delays in Qualified processes. Apply contingency margins (RT ASIC) schedule. in development schedule.

General 3B Increase in estimated Cost/Schedule Complex ASIC development. Standard ASIC flow. Good heritage and Apply margins in cost Purpose cost , lack of funding, /Programmatics experience in many EU budget to avoid overrun. Interface ASIC or possible delays in players. Apply contingency margins (RT ASIC) schedule. in development schedule.

DCM 3C Increase in estimated Cost/Schedule TRL2 in 2009. Very high cost Need for involvement of a Invest in technology and Apply contingency margins (Distribution cost , lack of funding, /Programmatics estimate. Expensive DSM major player in consortium testing in development schedule. and Command or possible delays in process. Long prototyping time with IP cores developers. Apply margins in cost Module-SoC schedule. of at least 4 years. Large budget to avoid overrun. ASIC) amount of software and auxiliary EGSE development. May require update of DARE (Design Against Radiation Effects) library and qualification of CCGA (Ceramic Column Grid Array) packages.

DCM 3C Critical technology Technical Complex prototyping, large Need for involvement of a Invest in technology and (Distribution issues affecting amount of software and major player in consortium testing and Command technical performance. auxiliary EGSE development with IP cores developers. Module-SoC required. May require update ASIC) of DARE (Design Against Radiation Effects) library and qualification of CCGA (Ceramic Column Grid Array) packages.

DCM 3B Critical technology Technical EMC/EMI issues. Full Appropriate EMC/EMI Invest in technology and (Distribution issues affecting qualification needed. Crowded testing to avoid or testing and Command technical performance. PCB. minimize risk. Module) Board

Generic 2B Increase in estimated Cost/Schedule Complex and costly module Invest in technology and Regarded as upgrade Apply margins in cost Thermal I/F cost , lack of funding, /Programmatics development/upgrade. testing. rather than new budget to avoid overrun. or possible delays in Estimated non recurring cost. development. Apply contingency margins schedule. in development schedule.

DHS (General) 2B Increase in estimated Cost/Schedule Availability of Power ST Power MOSFET for 60 Availability outside Europe: cost , lack of funding, /Programmatics MOSFETS. V and 100 V are in - USA: IR, International or possible delays in development & evaluation Rectifier (ITAR schedule. with ESA & CNES (valid implications) for 60 kRad, and maybe - Japan: Fuji + JAXA more with newer devices).

Table 16-7: Top Risk Log DHS Modules

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 MECHANISMS S/A 2B Increase in estimated Cost/Schedule/Progra Cost and schedule risk are Infineon is also developing Hinge heritage from The technology Deployment cost , lack of funding, mmatics considered moderate by CDF & evaluating a Power successful space development task for Mechanism or possible delays in mechanisms specialist. MOSFET with DLR. experience of French Nanosat is to scale down (SDM) schedule. Technology at TRL 4 in 2008, Microsat programme the MYRIADE hinge design expected to reach TRL 8 in 4 MYRIADE. in order to serve the task of years. deplying a max. S/A size of 34x23cm.

S/A 2A Critical technology Technical Dynamic performance during Low visibility on this Appropriate testing. Invest in technology. Deployment issues affecting launch (vibration amplitudes of development for ESA for Mechanism technical performance. flexible elements) and their the moment. (SDM) fatigue behaviour.

Deorbit 2B Increase in estimated Cost/Schedule Cost and schedule risk are Similar concepts (only Invest in technology and Apply contingency margins Deployment cost , lack of funding, /Programmatics considered moderate by CDF larger) are currently testing. in development schedule Mechanism or possible delays in mechanisms specialist. developed by RUAG (DDM) schedule. Complex system includes sail Aerospace Austria (RAA) and structure. Technology at for the GAIA sunshield. TRL 4 in 2008, expected to reach TRL 8 in 4 years. DDM is US technology and may be subject to ITAR.

Deorbit 3B Critical technology Technical Dynamic performance during Similar concepts (only Invest in technology and Deployment issues affecting launch (vibration amplitudes of larger) are currently testing. Mechanism technical performance. flexible elements) and their developed by RUAG (DDM) fatigue behaviour. Folding and Aerospace Austria (RAA) deployment of MLI. for the GAIA sunshield.

HDRM 3B Possible delays in Cost/Schedule The Frangibolt from TiNi Early start of ITAR TRL 8 in 2009 for US Invest in technology and schedule if US version /Programmatics Aerospace (USA) is subject to negotiations if US product product. Already existing testing of European HDRM. is selected. Increase in ITAR regulations. European is selected. product. estimated cost , lack of design is not available (TRL 2) funding, or possible delays in schedule for European version.

Nano- 3B Possible delays in Cost/Schedule US Product may be subject to Early start of ITAR TRL 7 in 2009 for US Invest in technology and terminator schedule if US version /Programmatics ITAR regulations. European negotiations if US product product. testing of European Nano- is selected. Increase in design is not available (TRL 2) is selected. terminator estimated cost , lack of funding, or possible delays in schedule for European version.

Table 16-8: Top Risk Log Mechanisms Modules

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 POWER Battery 2A Increase in estimated Cost/Schedule Uncertainties in budget or time Batteries are based on TRL 5 in 2009. Expected to Apply contingency margins cost , lack of funding, /Programmatics allocation for the required ABSL Li-Ion Battery reach TRL 8 by 2012. in development schedule. or possible delays in technology development. technology. Invest in Apply cost budget margins schedule. technology and testing. to avoid overrun.

Battery 3B Critical technology Technical No redundancy in battery cells Need for high quality/ high issues affecting (Nanosat policy). reliability components to technical performance. meet performance requirements.

Solar Array 2B Increase in estimated Cost/Schedule Currently very low TRL (TRL 1 Technology is based on Invest in technology and Apply contingency margins cost , lack of funding, /Programmatics in 2009). Thin film technology triple junction solar cell testing. in development schedule. or possible delays in development required. technology which is Apply cost budget margins schedule. already existing. to avoid overrun.

Solar Array 3B Critical technology Technical Criticality involved in the new Technology is based on Invest in technology and Need for high quality/ high issues affecting thin film technology. No triple junction solar cell testing. reliability components to technical performance. redundant solar array strings technology which is meet performance (Nanosat policy). already existing. requirements.

Power 2B Increase in estimated Cost/Schedule Uncertainties in budget or time Based on Buck (Step- Apply contingency margins Conditioning cost , lack of funding, /Programmatics allocation for the required down) DC DC converter in development schedule. Module (PCM) or possible delays in technology development. topology. Apply cost budget margins schedule. to avoid overrun.

Power 3B Critical technology Technical No redundancy. Need for high quality/ high Invest in technology and Conditioning issues affecting reliability components to testing. Module (PCM) technical performance. meet performance requirements.

Power 2B Increase in estimated Cost/Schedule TRL 4 in 2009. Expected to reach TRL 8 Conditioning cost , lack of funding, /Programmatics by 2011. Module (PCM) or possible delays in schedule. Table 16-9: Top Risk Log Power Modules

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 PROPULSION Propellant 1B Increase in estimated Cost/Schedule Custom tanks need TRL 5 in 2008. Expected Invest in technology and Titanium propellant tanks Tank cost , lack of funding, /Programmatics developing and qualifying. If to reach TRL 8 by 2012 testing. Apply contingency have been widely used for or possible delays in new material is used for under adequate funding. margings in development spacecraft applications. schedule. tank(depending on the schedule. Minor technical risk. propellant) the TRL timetable will change.Valves are COTS but may require to be developed and made smaller. Required weld qualification for new materials thicknesses, smaller PMDs for hydrazine tanks and minimistation of tank mounting mass.

Cold Gas 2B Increase in estimated Cost/Schedule The 'nanosat' version has a TRL 4 in 2008. Expected A larger 'microsat' sized Invest in technology and Generator cost , lack of funding, /Programmatics demonstration model, but no to reach TRL 8 in 2013. version has been space- testing. or possible delays in plans for qualification. The qualified and a smaller schedule. main issues for the nanosat 'cubesat' version is cold gas generator could be expected to be qualified in identified as development and 2009. qualification.

Solid 3B Increase in estimated Cost/Schedule Significant development and Invest in technology and Low development cost. Apply contingency marging Propellant cost, lack of funding, or /Programmatics qualification is required to testing. in development schedule. Thruster possible delays in prove the validity of fuels for schedule. nano space applications. TRL 2 in 2009. Closest example of appropriately sized motors is in hobby rocketry.

Solid 3B Critical technology Technical Uncertainties in the technical Invest in technology and Low development cost. Propellant issues affecting validity of hobby rocketry testing. Thruster technical performance. motors for nano space applications. Single 2B Increase in estimated Cost/Schedule TRL 3 in 2009. Major Expected to reach TRL 8 Most components being It is suggested also that the Thruster cost , lack of funding, /Programmatics development required in the by 2015. Apply put forward for this module thruster and valve Monopropellan or possible delays in electronics. contingency margins in have already been used in technologies could be t schedule. development schedule space applications. 'MEMSified' in order to (Thruster and latch valves reduce their currently available OTS TRL 9) significant mass.

Single 2B Critical technology Technical Technical uncertainties in the Invest in technology and Verify compliance with the Thruster issues affecting electronics. testing. 40Krad radiation tolerance Monopropellan technical performance. requirement (lifetime). t Single 2C Increase in estimated Cost/Schedule TRL 3 in 2009. Major Expected to reach TRL 8 Some of the components Invest in technology and Thruster cost , lack of funding, /Programmatics development required in the by 2015. currently being put forward testing. Apply contingency Butane or possible delays in electronics. for this module have marging in development schedule. already been used in space schedule. Apply cost applications. budget margins to avoid overrun.

Single 2C Critical technology Technical Technical uncertainties in the Invest in technology and Verify compliance with the Thruster issues affecting electronics. testing. 40Krad radiation tolerance Butane technical performance. requirement (lifetime).

Three Thruster 2C Increase in estimated Cost/Schedule TRL 3 in 2009. Major Expected to reach TRL 8 Some of the components Invest in technology and Butane cost , lack of funding, /Programmatics development required in the by 2015. currently being put forward testing. Apply contingency or possible delays in electronics. for this module have margin in development schedule. already been used in space schedule. Apply cost applications. budget margins to avoid overrun.

Three Thruster 2C Critical technology Technical Technical uncertainties in the Invest in technology and Verify compliance with the Butane issues affecting electronics. testing. 40Krad tolerance technical performance. requirement (lifetime).

Four Thruster 2A Increase in estimated Cost/Schedule Uncertainties in budget or time Technology is already at Flight model of microprop Apply contingency marging MEMS cost , lack of funding, /Programmatics allocation for the required TRL 5 in 2009. Expected system employing MEMS in development schedule. Nitrogen or possible delays in technology development. to reach TRL 8 by 2013. thruster is to be flown Apply cost budget margins schedule. aboard Prisma satellites in to avoid overrun. 2009. However, this is not a formal qualification of the components.

Four Thruster 3B Critical technology Technical Critical MEMS thruster pod Attention must be given to Invest in technology and MEMS issues affecting and MEMS integration. the MEMS thruster pod testing. Nitrogen technical performance. and the MEMS integration.

Table 16-10: Top Risk Log Propulsion Modules

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Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 UCTURES/HARNESS Conventional 1A Increase in estimated Cost/Schedule Delays during manufacturing, TRL 7 in 2009. Could No technology No technical issues. Multifunctional cost, lack of funding, or /Programmatics assembly and/or testing of achieve TRL 9 in 2 years. development issues. Structure possible delays in multifunctional structure. schedule. Innovative 3B Increase in estimated Cost/Schedule TRL 2 in 2009. Expected to Major investment and Multifuctional cost, lack of funding, or /Programmatics reach TRL 8 by 2016. technology development Structure possible delays in required to bring schedule. technology to required TRL level by 2018. Harness based 3C Non compliance with Cost/Schedule TRL 1 in 2008. Expected to Major investment and Verify compliance with the on Nanotubes Nanosat technology /Programmatics reach TRL 6 by 2017-2018. technology development 40Krad radiation tolerance readiness required to bring requirement (lifetime). requirements. technology to required TRL level by 2018. Conventional 1A Increase in estimated Cost/Schedule Delays during manufacturing TRL 7 in 2009. Could No technology No technical issues. Harness cost (lack of funding) or /Programmatics and/or assembly of achieve TRL 9 in 2 years. development issues. possible delays in conventional harness. schedule. Nano-D 2A Increase in estimated Cost/Schedule Current TRL is 4.The Validation of the Nano-D Or on the contrary adapt Space evaluation of the Connector cost, lack of funding, or /Programmatics Spacewire cable can be used connector for the the Spacewire cable to fit Nano-D connector will be possible delays in with the Micro-D connector but Spacewire cable. with the Nano-D. addressed by TEC-QTC. schedule. it is not sure that it will be TRL should reach 6 in 2-3 possible to use it with the years. Nano-D. Micro-D 1A Increase in estimated Cost/Schedule Delays during manufacturing Micro-D connector is Connector cost, lack of funding, or /Programmatics and/or assembly. available now and has possible delays in been validated for schedule. Spacewire (e.g. from Axon). Table 16-11: Top Risk Log Structures/Harness Modules

Subsystem / Risk Classification Module Index Risk scenario (Severity-Impact) Cause Mitigating Action 1 Mitigating Action 2 Mitigating Action 3 THERMAL Passive thermal control may not be sufficient to satisfy all Critical technology platform and payload needs in issues affecting the different mission Very low technical risk of Implement active thermal Black Paint 1A technical performance. Technical scenarios. passive thermal control. control if required. No programmatic issues.

Increase in estimated Apply contingency marging cost, lack of funding, or Item currently available in schedule. Apply cost Thin Plate possible delays in Cost/Schedule/Progra US developed module, may be from US vendor (Starsys Early start of ITAR budget margins to avoid Heat Switch 2B schedule. mmatics subject to ITAR. Research). negotiations, if applicable. overrun. MiSER Miniature Satellite Increase in estimated Energy cost, lack of funding, or Item currently available Regulating possible delays in Cost/Schedule/Progra US developed module, may be from US vendor (Starsys Early start of ITAR No technology Radiator 2B schedule. mmatics subject to ITAR. Research). negotiations, if applicable. development required.

Need for high quality/ high Critical technology reliability components to Kapton Heaters are issues affecting No redundancy implications meet performance currently available from Include redundant heater Kapton Heater 3B technical performance. Technical (Nanosat Policy) requirements. MINCO Inc. (Action taken). Polyimide flexible thermofoil heaters or Increase in estimated Kapton Heaters are cost, lack of funding, or currently available from MINCO Inc. has a possible delays in Delays during manufacturing, MINCO Inc.. No European representation No technology Kapton Heater 1A schedule. Cost/Schedule/Programtesting, and/or delivery. development required. office in France. development required.

T sensors (e.g. Resistance Increase in estimated Apply contingency marging Temperature cost, lack of funding, or Largely used in space in schedule. Apply cost Detectors possible delays in Delays during manufacturing, applications. NASA No technology budget margins to avoid (RTD)) 1A schedule. Cost/Schedule/Programtesting, and/or delivery. qualified. development required. overrun.

Some development will be Increase in estimated required for the miniturization Apply contingency marging cost, lack of funding, or and adaptation of heat pipes in development schedule. possible delays in for specific Nanosat Invest in technology and Apply cost budget margins Heat Pipes 2B schedule. Cost/Schedule/Programapplications. testing to avoid overrun. Multi-layer No technology Insulation development or (MLI) technical issues. Table 16-12: Top Risk Log Thermal Modules

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16.8 Risk Log Conclusions • Apply margins in schedule and budget to mitigate risk • 2018 target is in line with most module development time schedules. However, development time assumptions have been more optimistic than those described in the “European Space Technology Master Plan” • Heavy investment in technology and testing is required for modules with TRL 1-2-3 in 2009 and for ASIC development • The 40 krad radiation tolerance requirement will have an important impact on the cost and development time of the avionic system. Mission lifetime and environment (mission type) should be balanced with the cost and schedule limitations • High reliability components and short mission duration is a must in order to achieve desired performance without failure tolerance.

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17 PROGRAMMATICS This chapter describes the programmatic goals of this study, the requirements and design drivers which are particularly important in achieving the programmatic goals and the assumptions and trade offs made in this context. Accordingly, a programmatic approach is elaborated leading to a proposal for model philosophy, spacecraft level test matrix and a schedule estimate. 17.1 Requirements and Design Drivers In the scope of this particular study, “module” refers to a group of individual subsystems units, which, when combined, are able to offer a particular service(s) to the spacecraft, and which, for programmatics and costing purposes, can be treated as a single identifiable element. Therefore, the hierarchy in the product tree would be: System > Subsystem > Module > Unit > Component Each subsystem in the NanoSat comes as a set of complete, highly integrated, interchangeable modules. Thus, the system has a high modularity. The anticipated launch date of 2018 is likely to be an important design driver. 17.2 Assumptions and Trade-Offs For the preparation of the programmatic outputs, a highly modular and integrated system has been considered. The duration of the phase B in the schedule is a typical number, without consideration of specific NanoSat features. It is assumed that at least TRL 5 is achieved for all the modules at start of phase C/D. This start has been considered as January 2011. Depending on the chosen technology for each subsystem, this could be earlier or later (for example, if harness based on nanotubes is selected, the start would be later, as TRL 5 would be achieved only around 2015). This study does not take under consideration any launcher separation mechanism. The payload and the mission scenario have not been identified. Nevertheless, the payload is considered to be also modular. 17.3 Technology Development The Technology Readiness Levels (TRL) present a systematic measure, supporting the assessments of the maturity of a technology of interest and enabling a consistent comparison in terms of development status between different technologies. The different TRLs (as used by ESA and NASA) are defined in Table 17-1:

TRL 1 Basic principles observed and reported TRL 2 Technology concept and/or application formulated TRL 3 Analytical and experimental critical function and/or characteristic proof-of concept

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TRL 4 Component and/or breadboard validation in laboratory environment TRL 5 Component and/or breadboard validation in relevant environment TRL 6 System/subsystem model or prototype demonstration in a relevant environment (ground or space) TRL 7 System prototype demonstration in a space environment TRL 8 Actual system completed and “flight qualified” through test and demonstration (ground or space) TRL 9 Actual system “flight proven” through successful mission operations Table 17-1: TRL scale Table 17-2 shows an indication of the development time depending on the current TRL. According to the European Space Technology Master Plan, to prepare the contractual basis for multi-annual programs it takes about 18 months to reach political agreement on financial ceiling. This has also been included in the table.

TRL Duration 5-6 4 years + 1.5 year 4-5 6 years + 1.5 year 3-4 8 years + 1.5 year 2-3 10 years + 1.5 year 1-2 12 years + 1.5 year Table 17-2: TRL - development duration The development of the NanoSat is driven by the TRL of its subsystems. It is important to note here that not all the specialists within the NanoSat team have proposed developments for the modules for their respective subsystems according to Table 17-2, in that they have been more optimistic. 17.4 Model Philosophy 17.4.1 Case 1: LEO This section first points out the main characteristics, advantages and disadvantages of the prototype and protoflight approaches for the model philosophy. Second, it describes the selected approach for the NanoSat study, the hybrid protoflight philosophy. Prototype Philosophy [ECSS-E-10-02A (B.1.2.1)] The prototype approach is applied to projects with new and/or complex design and with special mission requirements. The advantages of this approach are: • Low risk

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• Possibility to perform parallel activities on different models • Completion of qualification activities prior to acceptance • Possibility to use EQM units as integration spare. The disadvantage is its high cost. Protoflight Philosophy [ECSS-E-10-02A (B.1.2.2)] The protoflight approach is applied to projects in which: • No critical technology is employed in the design • Qualified hardware is extensively used • Compromise is permitted to reduce cost, accepting a moderate level of risk. The pure protoflight approach is based on a single model (PFM) to be flown after it has been subjected to a protoflight qualification and acceptance test campaign (see ECSS-E-10-03 for details). The advantage of this philosophy is its low cost. The disadvantages are: • Increased risks • Serial of activity flow on the same model, leading to an increased time expenditure • Contextual thorough qualification and acceptance activities, increasing the time expenditure • No integration spares; no training models. The protoflight philosophy is the one usually followed by the “standard” nanosatellites developed at universities, research institutes, etc, mainly because it allows reducing costs with a moderate risk. In the case of the NanoSat object of this study, a great amount of new design is involved, which would suggest to follow a prototype approach. Therefore, looking at the previously presented philosophies, it can be concluded that the ECSS standard protoflight philosophy with one PFM only is on high risk, and so a hybrid protoflight philosophy can be considered: Hybrid Protoflight Philosophy [ECSS-E-10-02A (B.1.2.3)] • Compromise between prototype and protoflight • Advanced qualification activities in areas of new design or areas having a critical impact on the verification programme • Protoflight model is flown after a protoflight test campaign whose scope was reduced compared to that of a pure protoflight campaign • Specific qualification tests on dedicated models. The proposed model philosophy for the NanoSat consists of the following models: • Avionics Test Bench (ATB) • Protoflight Model (PFM)

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• Recurrent Flight Model (FM). Avionics Test Bench The ATB allows the verification of the proper interaction of all subsystems. For this aim, breadboard models of the different modules can be sufficient. Protoflight Model It is not foreseen to develop a structural-thermal model (STM). The NanoSat is small and it is highly integrated. This means its centre of gravity does not change much and no extreme thermal gradients are expected either. All flight hardware will be tested at system level with the protoflight model. The PFM shall ensure the first integration and verification of the structure, equipment and payload at full system level. Recurrent Flight Model After the first flight model has been developed, the next NanoSat will follow a recurrent spacecraft approach. A schedule for this second spacecraft is proposed to give an indication of the minimum time needed for a recurrent spacecraft. 17.4.2 Case 2: GTO From the programmatics point of view, the two different cases have no impact in the proposed model philosophy. Therefore, the model philosophy presented for the LEO case is applicable to the GTO case. 17.5 Options No options are identified which could impact the principal programmatics and AIV approach. 17.6 Spacecraft Level Test Matrix Table 17-3 shows the proposed tests to be performed on the protoflight model and the subsequent flight models of the NanoSat, together with the verification strategies for each test and model.

Test description PFM FM Mechanical Interface R, T R, T Mass Property A, T T Electrical Performance T T Functional Test T T Propulsion Test T T Thruster Lifetime Test T A Deployment Test A, T A, T

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Telecommunication Link T A, T Alignment A, T T Strength Load A, T T Shock/Separation T 1) T Sine Vibration A, T T Modal Survey A, T Acoustic A, T T Outgassing A, I I (T) Thermal balance A, T T Thermal vacuum (T) T Microvibration A, T T Grounding/Bonding R, T 2) R, T Radiation testing T T 3) EMC conductive interference T T EMC radiative interference T T DC magnetic T T RF testing T T Abbreviations: I=Inspection, A=Analysis, R=Review, T=Test Table 17-3: NanoSat test matrix Comments: 1) Shock/separation tests could be verified by electrical / command-wise check 2) Grounding/Bonding test on PFM depends on PFM design 3) Radiation test depends on instrument sensitivity. 17.6.1 Environmental Tests The environmental tests as requested above shall allow demonstrating that the satellite operates nominally under environmental conditions corresponding to the worst cases of the mission. Static load/Strength load test: This test is required to demonstrate that the structure is able to support the static load constraints corresponding to the design load and is considered as part of the qualification sequence. Modal survey test: This test is considered as part of the qualification sequence and performed on the PFM only. It shall be combined with the sine vibration tests, as it consists in particular measurement performed during the low level tests using specific test instrumentation. Sine and acoustic vibration tests: These tests are considered as part of the qualification sequence. As they are performed on the PFM, these tests shall also cover the acceptance objectives.

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Separation and shock test: This test has been considered from the mechanical point of view only and should be performed on the PFM as part of the mechanical qualification. It’s foreseen to repeat such a test at PFM level with the satellite in electrical operational configuration (potential impact of the shock on electronic units, relays and electrical contacts - potentiality to provoke spurious signals or transient spikes). Electromagnetic compatibility tests: The measurements to be performed pending on the EMC specification, during this test phase an adapted System Functional Test (SFT) should be performed. DC Magnetic measurement test: This test shall be performed on system level according to the experiment, mission requirements. The total DC magnetic field generated by the satellite shall be predicted and or tested (from analysis of the test data) at the magnetometer location to be compared with the specification. Compensation at system level is foreseen if needed. Thermal balance test: The thermal balance test will verify the thermal model and the thermal control efficiency and performance. This test should be performed with the PFM in a vacuum chamber with solar simulation. This complete thermal balance test shall simulate several significant mission cases. As it is performed on the PFM, it allows the verification with the electronic units and to verify the active thermal control operating in flight representative configuration. During this test phase the full scientific payload test (adapted SFT/SPT) could be performed. The thermal cycling test could continue the thermal balance test without change of configuration neither of the chamber nor of the satellite or the adapter installation. The thermal test adapter shall enable the position of the spacecraft configuration with respect to the solar simulator to be selected during the thermal test such that it complies with the test requirements. Thermal vacuum (cycling) test: The thermal cycling test is proposed in the PFM sequence as part of the acceptance of the satellite in addition to the limited thermal balance test. It would enable the verification of the satellite operation in cold and hot cases corresponding to the expected limit temperatures. For these two cases the adapted SFT should be repeated. 17.7 Subsystem TRL The different subsystem specialists have prepared tables containing information about the development schedule for each of the modules proposed in the study. This schedule is presented in Table 17-4, showing also the TRLs. The specialists also provided inputs for the development activities. These activities have been grouped under their corresponding entry in the table, for ease of reading, but can be seen by double-clicking on the table. Note on DHS modules: Not all the required inputs for the programmatic analysis of this subsystem were provided by the DHS expert.

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DHS Control Distribution Unit 2489 General purpose Interface ASIC 358 DCM (SoC ASIC) -System on a chip-

Power Solar Array 23456 Battery Pack 23456789 Power Conditioning 45678

Propulsion Propellant Tank Module (A) 56778 Cold Gas Generator Module (B) 456678 Solid Propellant Thruster Module (C) 2334556678 Single Thruster Monopropellant Module (D) 34556678 Single Thruster Butane Module (E) 34556678 Three Thruster Butane Module (F) 34556678 Four Thruster MEMS Nitrogen Module (G) 34556678

Structure Conventional Structure 79 Innovative Structure 2 3456789 Harness based on Nanotubes 12456 Conventional Harness 79

Mechanisms S/A Deployment Mechanism (SDM) 456 8 Deorbit Deployment Mechanism (DDM) 456 8 Hold down and Release Mechanism (HDRM) 28 Nano-Terminator Deorbit Module (NTDM) 28

Thermal Black paint 8 MiSER (Miniature Satellite Energy Regulating Radiator) 8 Thin Plate Heat Switch 8 Heater line (2 heaters+1 sensor) 8 Heat pipe 8 MLI blankets 8

Table 17-4: Subsystem Modules’ TRL

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17.8 Schedule A possible schedule for the NanoSat protoflight spacecraft and the first recurrent spacecraft has been developed and is shown in Figure 17-1.

Figure 17-1: NanoSat project schedule 17.9 Summary and Conclusions The anticipated schedule start is driven by the readiness of all the subsystems and the complexity of the chosen technology. As already mentioned in the assumptions section, at least TRL 5 has to be achieved for the start of phase C/D. The duration from start of phase C/D until the launch of the protoflight model is estimated to be nearly four years. It is driven by the procurement of EEE parts and by the complexity of subsystems and payload. The propulsion subsystem is not shown on the critical path, but if it comes as highly integrated block it might be schedulewise considered as the other PFM units in the schedule. The time to build a recurrent spacecraft up to launch is estimated to be about two years, provided that no extra time is needed for EEE parts procurement. The payload is also considered to be modular and available at start of the FM integration. The presented schedule is an engineering estimation based on experience. The real schedule for such a NanoSat spacecraft will be driven to a considerable extent by the payload, the envisaged mission and the complexity of the subsystems finally chosen.

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18 COST This chapter presents a description of the cost estimates of the Nanosat Individual Modules Development. The actual detailed estimates are presented in a separate cost report (CDF-84(B)).

The detailed estimate is based on the CDF design provided by all of the technical disciplines, according with Risk and Programmatics assumptions. For cost estimate purposes, three different LEO mission Options have been analysed for the Nanosat Platform and Operations, as follows (note these are different than the mission options presented in the Systems chapter): • Option 1: LEO Mission Scenario limited to the case of absence of the Propulsion subsystem. In this case the 1st NR Nanosat Platform and Operations cost estimates have been performed; • Option 2: LEO Mission Scenario limited to the case of absence of the Propulsion subsystem. In this case the 9th R Nanosat Platform and Operations cost estimates have been performed. The recurring approach has been based on the assumption of an identical HW and SW technology to the one proposed for Option 1. Due to the wide range of possible LEO mission configurations for the Nanosat Platform mission and in order to present all the possible cost limit cases, an additional cost assessment has been performed assuming a different Payload from Option 1. • Option 3: LEO Mission Scenario limited to the case of absence of the Propulsion subsystem. In this case the 9th NR Nanosat Platform and Operations cost estimates have been performed. The non recurring approach has been based on the assumption of an identical HW and different SW technology from the one proposed for Option 1. For all the LEO Mission Options it has been assumed that the Project will be implemented by a small and low-cost Prime that may propose `light` system level procurement approach and solutions. 18.1 Class of Estimate The cost estimates have been performed within the CDF environment by ESA/ESTEC Cost Engineering (TEC-SYC). The type of estimate for is Class 4 to 5 (Class 4 in general, but Class 5 for Operations and ESA Internal Costs), as described in the ESA Cost Engineering Chart of Services RD[30]. The project is considered to be of Major Complexity and the accuracy of the total estimate is expected to be about 25%. 18.2 Cost Estimate Methodology The following methods have been used, in descending order of preference: • Reference to similar equipment/subsystem/system level costs, taking into account the amount of new development required • Expert judgement from technical specialists in combination with cost references, in case the amount of new development is extensive • Expert judgement from technical specialists only, if references are not available • Equipment cost models

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• System-level cost relationships (for the Prime contractor and subcontractor activities), based on observed relationships for relevant references • The ESA internal, subsystem level cost model RACE in combination with expert feedbacks from the specialists involved in the Study. 18.3 Scope of Estimate 18.3.1 Individual Modules The detailed cost estimates for the Nanosat Modules Development include: • Individual Modules Predevelopment Provision • Individual Modules Development Programme Coordination Provision • Individual Modules & Software NR Cost up to the 1st PFM • Learning Analysis • Individual Modules Cost Risk margin. 18.3.2 LEO Mission Scenario (Option 1, 2, and 3) For all the three Options the detailed cost estimates for the Nanosat Spacecraft Platform include: • Spacecraft Platform Phase B ROM • Spacecraft Platform Phase C/D Industrial development and production cost • Spacecraft Platform Phase C/D Industrial System Level Management & Control, Engineering, Product Assurance, Ground Support Equipment and AIT/AIV cost • Spacecraft Platform Phase E1 of the mission ROM • Operations ROM (provided by ESOC, including Risk) • Spacecraft Platform Industry and ESA Internal Cost Risk Margins. Not included in the estimates are: • Payload/Instrument costs • Launch costs • Phase A costs, which are assumed to be covered under the ESA Technology Development Programs (GSP, TRP, GSTP). 18.4 Main Assumptions The various requirements and assumptions described in the basic study documentation apply to the cost estimates. In addition, the specific cost-related assumptions described hereunder were made. 18.4.1 General Cost Assumptions • All technical details relevant to the various modules have been extracted from the CDF Nanosat Workbooks • TRLs today and their evolution have been assessed by experts, as included in the Modules (Pre) development planning assumed by Programmatics

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• Model Philosophy have been assessed by experts, as included in the Modules (Pre) development planning assumed by Programmatics • The System Level estimates for the LEO Mission have been based on a Low Cost Procurement approach by the possible Platform Prime Contactor (light PA and management, simple industrial set-up of the Project). 18.4.2 Technology Development Assumptions Technology developments are assumed to be performed before start of a mission Phase C/D in order to reach a sufficiently high TRL level for various critical technologies assumed to be developed under ESA Technology Research and Development activities. All Pre development provisions include: • ASIC Predevelopments for the electronics and interfaces • EGSE/SCOE/Tooling/Overheads necessary for each module • Radiation hardening and tests to withstand 40 kRad • Miniaturization activity according with TRL schedule provided by the technical specialists • The production of all required interface documentation • Nanosat Modules Development Programme Coordination activity • ITAR free additional provision for the identified modules. Pre development provisions have been reported for the following Nanosat modules: • Digital Sun Sensor • Star Tracker • Reaction Wheel • GNSS receiver • Magnetometer • Magnetic Torque • Navigation Camera • Coarse Gyro • Module A Integration • Tank Assembly (B) • Module B Electronics • Module B Integration • Thrusters(C) • Module C Electronic • Module C Integration • Module D Electronics • Module D Integration • Module E Electronic

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• Module E Integration • Module G Electronic • Module G Integration • Module F Integration • Battery • Solar Array • Converter Module • Harness based on Nanotubes • Mobile Phone based Transponder • Multifunctional Distributed Ant • Electronics for antenna systems • Lightweight X band antenna • S-band UP/DW Converter Efficient Power Amp. (2) • RT ASIC Interface • SoC ASIC (OBC) • Generic Thermal Interface • Fully integrated Panel Bottom (B) (1) • Fully integrated Panel Top (B) (1) • Fully integrated SA Panel (B) (1) • Half Wall Panel (B) (1) • Panel Bracket (B) • SDM (S/A Deployment Mechanism) • DDM (Deorbit Deployment Mechanism) • Nanoterminator • HDDM (Hold Down and Deployment Mechanism) • On-Board AOCS Software. 18.4.3 Modules and LEO Mission Schedule Assumptions • Modules TRL’s today and their evolution have been indicated by experts, as included in the Modules (pre) development planning assumed by Programmatics. • The LEO Mission Scenario cost estimates here presented have been based on the schedule provided by Programmatics (see Figure 17-1). 18.4.4 Ground Segment and Operations Cost Assumptions According with ESOC, an average mission lifetime of 3 years has been assumed for the LEO Options. Furthermore, it is necessary to highlight that the costs could have a high variability as they are strongly related to the mission rather than the spacecraft. For the routine provision it has been assumed: • Low cost, non guaranteed service

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• No redundancy in the ground system • No manpower outside working hours • No mission planning. Preparation costs have been based on a standard/average schedule of the LEOP and commissioning phase. 18.4.5 ESA Internal Costs Assumptions ESA Internal Costs have been estimated based on average ESA internal cost percentages for past IOD missions. This estimate should be considered as a ROM number to be verified by detailed planning within the project team. 18.5 Technology Readiness Level Definition The Technology Readiness Levels (TRL) presents a systematic measure, supporting the assessments of the availability and maturity of a technology of interest and enabling a consistent comparison in terms of development status between different technologies. The different levels as used currently in ESA as defined in an internal working group based on NASA’s Technology Readiness Levels are given in the table below:

TRL 1 Basic principles observed and reported TRL 2 Technology concept and/or application formulated TRL 3 Analytical and experimental critical function and/or characteristic proof of concept TRL 4 Component and/or breadboard validation in laboratory environment TRL 5 Component and/or breadboard validation in relevant environment TRL 6 System/subsystem model or prototype demonstration in a relevant environment TRL 7 System prototype demonstration in a space environment TRL 8 Actual system completed and “flight qualified” through test and demonstration TRL 9 Actual system “flight proven” through successful mission operations Table 18-1: Definition of Technology Readiness Levels 18.6 Cost Risk/Opportunity 18.6.1 Definition and Background A cost risk analysis has been performed by employing triangular cost distributions (Minimum, Most Likely, Maximum) and Monte Carlo-like simulations. As the Most Likely numbers, the point cost estimates as presented in the detailed estimate tables have been taken. The employed spreads from Minimum to Maximum take into account the uncertainties in the cost estimate relationships, quality of the cost model input parameters, quality and applicability of the references and cost estimate relationships used, and the possible variations in the amount of equipment modifications and qualifications required.

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All cost items in the estimate have been correlated amongst each other (i.e. the higher cost of one item increases the chance of a cost increase in the other items as well). The Cost Risk Margin has been established for a 70% confidence level (i.e. the chance that the budget including the Cost Risk Margin is sufficient for the project is 70%, or in other words the chance of a cost overrun is 30%). This Cost Risk Margin consists of several components: • Design Maturity Margin (DMM), to account for additional costs caused by unseen complexities that will be revealed as the design gets into more details. This entropic effect is inherent to the design process and therefore has to be provisioned as part of the core estimate. It is allocated 100% to Industry. • Cost Model Accuracy (CMA), to account for uncertainties in the cost estimates. It includes the contribution of the Inherent Quality of the cost Models (IQM) together with contextual factors such as the Degree of Adequacy (DOA) of the cost models used with respect to the specific context of the cost estimate, and the Quality of the Input Values (QIV). Assuming that industry has better and more detailed cost models than ESA because based on internal costs, 25% of the CMA is accounted for industry and 75% for ESA. • Project Owned Events (POE), to account for cost risks induced by potential negative events, as well as potential cost reduction opportunities, that may occur or not and that are under the direct responsibility of the Project Manager. POE risks are subject to mitigation measures to be managed at Project level. Assuming that industry has better and more detailed cost models than ESA because based on internal costs, 25% of the POE is accounted for industry and 75% for ESA. • External to Project Events (EPE), to account for cost risks or opportunities that originate from external influences out of the direct control and responsibility of the Project Manager. The EPE should normally belong 100% to ESA, but ESA regularly transfers the coverage for fair Geo-Return cost impact to Industry. The EPEs included in this estimate are preliminary provisions only and account for impacts of abnormal geo-return constraints, funding delays, etc. Depending on the actual constraints put on the program to reach implementation, this amount may increase compared to the value quoted in this cost report. The various elements of the total Cost Risk Margin are distributed over the Industry and ESA part of the costs. The costs for the Operations are not taken into account as part of the Industrial costs, therefore the risk and opportunities associated are fully assigned to the ESA part of the Cost Risk Margin. 18.6.2 Cost Risk/Opportunity Specific Assumptions 18.6.2.1 Individual Modules Specific Assumptions The following specific cost risks and opportunities have been identified: • CMA o IQM: The inherent quality of the cost models is considered as high for all modules which costs have been estimated via parametric methodology and when the TEC-SYC

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models have been verified to be applicable and well calibrated. A higher IQM risk margin has been applied (medium) when other estimating techniques have been used. o DOA: All cost items with a high IQM have been considered to have been set on a highly adequate model. Cost items with a medium IQM margin and with a good expert level of confidence/costing feedback have been set with a set on a highly adequate model as well. Cost items with a medium IQM margin and with a partial uncertainty expert level of confidence/costing feedback have been set on a medium DOA. In order to provide a consistent pre development provision for all the ASIC based modules, the DOA Type has been set on `Bias +, for Reaction Wheels and Gyro` when the output is probably overestimating, on `Bias -, for Magnetic Torque Module` when the output probably underestimating. In all the other cases an unknown biased DOA has been assumed. o QIV: the various inputs for the cost estimates at module level have been provided directly by the technical expert and are thus considered as of high value. Some exceptions have been identified and set on `medium` because a more consistent risk margin has been foreseen for specific ongoing technology development research activities costs and related experts feedbacks on them. • DMM: at module level, a wide range of Design Maturity Margin percentages have been identified, according with the independent Modules Pre development planning and TRL evolution in time. A 20% of DMM has been assumed for all the modules with very low TRLs, 15% for those that have been reached a technology readiness close to 5, and 5- 10% for all the demonstrated technology cost items. • POE and EPE: in addition to standard POE and EPE risk margin at equipment and system level (schedule slippage, geo-return competition…), specific risk/opportunities items have been taken into account for the Nanosat modules cost estimate according and integrated with the Modules Development Risk assessment : The following specific POE cost risks and opportunities have been identified: • Schedule Delays due to Predevelopments identified difficult identified technology issues and qualification, with a cost estimated to be about 5% of the total modules point estimate. The assigned probability is 60% and a 50%/50% cost split between Industry and ESA has been assumed. • Schedule Delays due to additional Development Models Required, with a cost estimated to be about 5% of the total modules point estimate. The assigned probability is 30% and a 50%/50% cost split between Industry and ESA has been assumed. • Additional design and qualification processes for the ASIC based modules (uncertainties on modules development process), with a cost impact of 10% of the total ASIC based modules point estimate. The assigned probability is 50% and a 50%/50% cost split between Industry and ESA has been assumed. • Unidentified risk, with a cost estimated to be about 6% of the total modules point estimate. The assigned probability is 50% and a 0%/100% cost split between Industry and ESA has been assumed. • Unidentified opportunities, with a cost estimated to be about 2% of the total modules point estimate. The assigned probability is 50% and a 0%/100% cost split between Industry and ESA has been assumed.

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The following specific EPE cost risks and opportunities have been identified: • The cost estimate foresees a risk provision for fair geo-return requirements transferred to the Contractor, with a cost estimated to be about 5% of the total modules point estimate. The assigned probability is 60%, with a 100% share for Industry. • The cost estimate foresees a risk provision for more stringent geo-return requirements transferred to the Contractor, with a cost estimated to be about 6% of the total modules point estimate. The assigned probability is 30%, with a 100% share for ESA. • The cost estimate foresees a risk provision for Constraints to open competition, with a cost estimated to be about 10% of the total modules point estimate. The assigned probability is 70%, with a 100% share for ESA. • Unidentified opportunities, with a cost estimated to be about 2% of the total modules point estimate. The assigned probability is 30% and a 0%/100% cost split between Industry and ESA has been assumed. 18.6.2.2 LEO Mission Scenario and Options Specific Assumptions The following specific cost risks and opportunities have been identified: • CMA o IQM: Please refer to section 18.6.2.1 or explanations on the cost risk assessment methodology that has been applied to the LEO Mission Options. o DOA: Please refer to section 18.6.2.1 or explanations on the cost risk assessment methodology that has been applied to the LEO Mission Options DOA magnitude. An unknown biased DOA has been assumed for all the cases, except for system level activities costs as .Management, PA, and Engineering o QIV: Please refer to section 18.6.2.1 or explanations on the cost risk assessment methodology that has been applied to the LEO Mission Options. o DMM: Please refer to section 18.6.2.1 or explanations on the cost risk assessment methodology that has been applied to the LEO Mission Options. o POE and EPE (including register based cost risks and opportunities): Please refer to section 18.6.2.1 or explanations on the cost risk assessment methodology that has been applied to the LEO Mission Options. 18.7 Conclusions and Recommendations A cost estimate with sufficient accuracy for this stage is available for the Nanosat modules and three different options applicable to a LEO mission scenario that will carry the identified modules. The main conclusions and recommendations of the cost evaluation of the CDF Nanosat mission concept are: A heavy investment for the Development of the modules has been identified in the Predevelopment activities. The main cost drivers are: • ASIC electronic technology and interfaces • 40 kRad radiation tolerance requirement at component level • Nanosat Modules Integration and Coordination activity.

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It is important to highlight that the 1st NR PFM modules cost figure could be the main obstacle for potential Industrial customers interest on the development of such non standard products, where a particular effort is needed to meet the Nanosat challenges. As first approach, a module's cost reduction and break even point analysis has been simulated via learning effect assuming a total of 10 units produced for the possible LEO mission scenario. A better cost analysis could not be foreseen for such a detailed Nanosat Platform at this stage, because driven by: • A more detailed business plan baseline • A detailed Procurement approach • Specific Market Effect scenarios. A consistent Platform cost reduction has been foreseen only for the Option 2, in case of a totally recurrent HW, SW, and assuming an identical Payload to be carried. As direct consequence, a real `low cost` platform looks to be achievable only in case of highly recurrent modules (HW & SW) applications and configuration, which allow to minimize System Level costs as well. A high percentage of the estimated total cost (excluding ESA Internal costs) has been attributed to various cost risk margins, due to: • The complexity of the individual modules development under demanding requirements as high performances, modularity, and miniaturization • The detailed modules development process planning that has been assumed to be the one specified by Programmatics at the moment • The current limited available specific definition of the independent modules • The current not specific definition of the mission/projects with the associated wide range of uncertainties • Consistent political and technical constraints to open competition scenarios. It is highly recommended to maintain a competitive industrial scenario for: • The modules to be developed • The hypothetical future Projects that will carry the identified units. Furthermore, it is highlighted that the market effect will have a strong impact on the negotiation activities because of the already identified monopoly situation. 18.7.1 Price Growth Phenomenon and Proposed Containment Measures Recent past projects have shown extensive cost growth in comparison to their initial implementation phase prices. Initial estimates systematically have been overshot even though technical difficulties and risks were considered and cost models calibrated on previous projects. The main reason for price growth is due to the fact the growth in itself is a dynamic process driven by the interaction between the different actors of the project. It therefore can not be assumed that systematically heavier initial cost estimates will lead to stable prices. A path to solving the cost growth problem is therefore: • To perform an internal, non-biased and independent cost estimate, defining a robust and clear baseline

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• To receive from Industry a complete, fair and substantiated estimate in the conceptual phases. This can be enforced by expressing rigorous costing requirements as foreseen in ECSS M60B and verifying their applications • To analyze all causes that feed the dynamic process of the price growth • To conceive or adapt the management methods and tools that will allow containing it. The following elements are identified as causes currently aggravating the price growth dynamic: • Direct negotiations are known to bring low compliance to requirements and excessive prices, especially for prime contractors • There is also an incentive for industry not to consider challenging tight schedules in this depleted programming situation but rather maintain smooth workload over time for their resources. This leads to schedules that may be longer than necessary with the consequence of higher prices (“marching army” effect). Identified price growth containment measures: • Reinstate Cost-Plus Fixed Fee type of contract for prime contractors, eventually including cost saving incentive clauses • Define a realistic and complete work plan for pre-developments of low TRL (<4) elements compatible with the overall schedule • Develop Earned Value Management (EVM) system during the implementation phase. This should allow identifying discrepancies earlier in time (not waiting for formal reviews) and getting a proper measure of the work physical progress and contractors’ productivity. Physical progress requires the definition of clear milestones at detailed activities level. These detailed activities milestones must be defined and closely watched by the project control in order to develop an intimate knowledge of the work progress and a proper measure of the difficulties met. • Develop predictive capabilities of agency’s project control by improving Estimate To Completion (ETC) expertise. The early detection of discrepancies via the physical progress measurement must be followed by the rigorous analysis of the corrective measures to be put in place, and the cost/schedule quantification of their implementation. • Improve visibility and control of labour resources by reforming industrial rates structure. Limited generic set of rate categories per qualification level (managers, engineers, technicians and workers) rather than categories per cost centres would allow to trace companies’ productivity and avoid disruption in traceability when companies’ structures are changing • Ensure that sufficient funding is assigned to the definition phase (this has been recently done) with the verification that the funds are used to mobilize resources and perform activities that actually contribute at best to the design progress and the technical risks deflation • Ensure that all needed pre-developments and long lead items are identified, scheduled in time and benefit from adequate funding profile.

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19 CONCLUSIONS 19.1 Satisfaction of Requirements This chapter reports the findings of the assessment study on the NanoSat mission based on the Mission Objectives, the derived requirements and constraints as provided by the customer. This chapter highlights the overall and critical aspects resulting from the study performed by the CDF team, in addition to the discipline specific conclusions which have been reported in the previous chapters. 19.1.1 Study Objectives The NanoSat study provided a high-level assessment of the technical feasibility of a nanosatellite platform based on a modular approach. The conceptually designed reconfigurable platform, assembled from a ‘kit’ of recurring modules, was shown to be suitable for supporting a wide range of payloads in a wide range of mission scenarios. Additional possible scenarios of interest for such a platform were identified (e.g. lunar missions, asteroid exploration etc) but their technical feasibility could not be assessed in the frame of the Study. The CDF study has shown, at a preliminary level, that a high performance nanosatellite platform can be technically feasible within the next decade, if the required engineering effort and development funding is timely made available. The three study case mission scenarios investigated throughout the study showed that the modules considered and presented can be sufficient to cover requirements and requests related to a wide spectrum of possible applications and to feasibly provide sufficient onboard resources, mass and volume to a miniaturised payload in different typical application cases. The initial study objectives and requirements have been shown to be realistic targets, as shown on the following table.

Initial Target Demonstrated in Study Cases Overall platform mass <10 kg (Goal), <20 kg From 9.1 kg (without payload and with no (max) onboard propulsion or RCS modules) To 18.5 kg without payload (depending mainly on the power and ∆v needs) Payload Mass: 5 kg Assuming a total mass limit of 20 kg, then payloads up to 10.9 kg are possible depending on power and ∆v need. Available payload volumes up to 22 litres (with no onboard propulsion or RCS modules) Power to payload: 10W Continuous 10 W power to payload shown to be feasible in mission scenarios with no eclipse (50% duty cycling in LEO cases)

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Initial Target Demonstrated in Study Cases Data Storage:>32Gb Concept expected to support >>32Gb within this timeframe Dta Downlink: 2 Mb/s A conservative 2.5 Mb/s useful data rate baselined with possibility to increase ∆capability: TBD m/s Up to 110 m/s possible (60 m/s via cold gas, 50 m/s via solid fuel) (assuming a 20 kg spacecraft mass) Table 19-1: Initial targets compliance table Several key ideas and concepts have been instrumental in being able to demonstrate the requirements as being acheivable. The main ones are: • Adoption of a fully single string platform concept, with redundancy to be provided at mission level • Modular multifunctional structure: allows a large reduction in recurring cost, elimination of NRE for future missions and greatly aiding the overal modularity and adaptability of the platform. • Elimination of analogue and point to point interfaces: possibly allowing to largely reduce the harness mass while greatly easing modularity and allowing the support of plug and play support to the modules. • Re-partitioning of traditional subsystem interfaces: essential to allow the definition of completely recurring modules and allows optimisation of system mass. • Widespread use of high levels of on-chip integration using radiation hard techniques: essential to ensure reliability and radiation hardness targets are met as well as the miniaturisation goals. The use of standard IP cores reduces the development cost and lowers the risk of interface problems. 19.2 Final Considerations Based on the NanoSat modular design and development philosophy, the CDF study has assessed the technical feasibility of the overall platform concept and suitability to a wide set of mission requirements. The obtained results are however based on a disruptive approach at system level (single chain, system margin only, etc...) and require an extensive multi-year development programme involving significant investments. A heavy investment for the module development has in particular been highlighted for the predevelopment activities. A spacecraft manufacturing cost reduction is however expected for the production of successive fully recurrent platforms. Furthermore, it must be highlighted that the real schedule for a mission based on a NanoSat spacecraft will be driven to a considerable extent by payload, envisaged mission and complexity of subsystems finally chosen.

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20 REFERENCES RD[1] Fundamentals of Astrodynamics and Applications, D. A. Vallado, 2nd Edition, The Space Technology Library. RD[2] CDF Study Report: Sentinel 5 Precursor, CDF-80(A), ESA, Oct. 2008 RD[3] CDF Study Report: Vegetation-GF, CDF-65(A), ESA, Dec. 2007 RD[4] Space engineering Mechanical - Part 3: Mechanisms, ECSS-E30-3, latest issue, http://www.ecss.nl RD[5] Application of “MAEVA” Hinge to MYRIADE Microsatellites Deployment needs, Jacques Sicre et al, proceedings of the 11th European Space Mechanisms and Tribology Symposium (ESMATS), 2005, Switzerland RD[6] Datasheet “Frangibolt Actuator Model FC2-16-31SR2”, TiNi Aerospace Inc., USA, http://www.tiniaerospace.com/fbt/fbfc2-16-31sr2.html RD[7] Datasheet “Frangibolt Fasteners”, TiNi Aerospace Inc., USA, http://www.tiniaerospace.com/fbfasteners.html RD[8] Datasheet “nanoTerminator™”, Tehters Unlimited Inc., USA, http://www.tethers.com/SpecSheets/nanoTerminator.pdf RD[9] Datasheet “nanosat Release Mechanism”, Tethers Unlimited Inc., USA, http://www.tethers.com/SpecSheets/nanosat%20Release%20SpecSheet.pdf RD[10] Paper “Technology Demonstrator of a Standardized Deorbit Module Designed for CubeSat and RocketPod Applications”, Nestor R. Voronka et al, SSC05-XI-4, Proceedings of the 19th Annual AIAA/USU Conference on Small Satellites, 2005, USA RD[11] D4187 Diagram, Online Datasheet, Ardé, 2002, Available at http://www.ardeinc.com/Liquid%20Propellant%20Storage%20Vessels/d4187%20files/4 187.htm RD[12] Spacecraft Propulsion Valves, Online Resource, Astrium, 2009, Available at http://cs.astrium.eads.net/sp/SpacecraftPropulsion/Valves/Spacecraft_Propulsion_Valves. htm RD[13] Innovative Plug and Play Micro propulsion systems, Online Datasheet, TNO, 2009, Available at http://www.tno.nl/downloads/DV%2007d002.pdf RD[14] Development of Micropropulsion Technologies for Microsatellites in the Netherlands, Presentation Slides from the Space Propulsion Conference, Sanders et al, 2008. RD[15] Pressure Transducer, Datasheet 1560, Presens. RD[16] Highlights of Nanosatellite Propulsion Development Program at NASA-Goddard Space Flight Centre, SSC00-X-5: Paper submitted to 14th Annual/USU Conference on Small Satellites, Rhee et al, 2000, Available at http://www.gkllc.com/lit/misc/Goddard- Nanosatellite_Propulsion_of_MEMS_Hydrogen_Peroxide_Development.pdf

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RD[17] Monopropellant Hydrazine Thrusters, Online Resource, Astrium, 2009, Available at http://cs.astrium.eads.net/sp/SpacecraftPropulsion/MonopropellantThrusters.html#CHT0 5 RD[18] Spacecraft Fluid Controls, Product Catalogue, MOOG, version 2.5, 2006, Available at http://www.moog.com/media/1/SpacecraftFluidControlsCatalog06.pdf RD[19] Cold Gas Thruster Valve SV14, Datasheet, Marotta RD[20] Thermal Throttle TT01, Datasheet, Marotta RD[21] MEMS for Space, Information Sheet, NanoSpace RD[22] MEMS-based micropropulsion system, Online Resource, NanoSpace, 2007, Available at http://www.nanospace.se/?id=8200 RD[23] Low Profile Integrable Inductor Fabricated based on LTCC Technology for Microprocessor Power Delivery Applications, IEEE Xplore, M. H. F. Lim1, Z. Liang1, and J. D. van Wyk, 2006. RD[24] Space Mission Analysis and Design, Third Edition. James R. Wertz. RD[25] “Flexible serially concatenated convolutional turbo codes with near-Shannon bound performance for telemetry applications”, CCSDS 131.2-O-1, September 2007. RD[26] Space Project Management, Risk management, ECSS-M-ST-80C, 31 July 2008. RD[27] PA & Safety Strategy for In-Orbit Demonstration Lightsat Concept, D/TEC-Q Technical Note, 29 January 2008. RD[28] ECSS-E-10-02A, “European Cooperation for Space Standardization, Space Engineering, Verification”, 17 November 1998. RD[29] ECSS- E-10-03, “European Cooperation for Space Standardization, Space Engineering, Testing”, 15 February 2002. RD[30] ESA Cost Engineering Chart of Services RD[31] TEC-SYC/GRE/SA/2005/021 ESA Cost Risk Assessment Procedure RD[32] TEC-MMA/2008/074 The NEOMEX Strawman: Enabling a Microsystem-Based Nanospacecraft

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21 ACRONYMS

Acronym Definition ACM Adaptive Coding and Modulation ADC Analog to Digital Converter AGC Automatic Gain Control AIT/V Assembly, Integration and Test/Verification AIV Assembly, Integration and Verification AMR Anhysteretic Remnant Magnetisation AOCS Attitude and Orbit Control System APE Absolute Pointing Error API Application Program Interface APSK Amplitude and Phase Shift Keying ASIC Application-Specific Integrated Circuit ASW Central Software ATB Avionics Test Bench BB Bread Board BCDR Battery charge/Discharge Regulator BM-SA Body Mounted Solar Array BOL Beginning Of Life CAD Computer Aided Design CAN Controller Area network CCGA Ceramic Column Grid Array CCM Constant Coding and Modulation CCSDS Consultative Committee for Space Data Systems CDF Concurrent Design Facility CDMU Command and Data Management Unit CER Cost Estimation Relationship CFRP Carbon Fibre Reinforced Plastic CGG Cold Gas Generator CIL Critical Items List CL Current Limiter

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Acronym Definition CLCW Command Link Control Word CM Core Module CMA Cost Model Accuracy CNES Centre National d’Etudes Spatiales CoM Centre of Mass COMMS Communication Subsystem COTS Commercial Off-The-Shelf CPU Central Processing Unit DAC Digital to Analog Converter DARE Design Against Radiation Effects DC Direct Current DCM Distribution and Command Module DCM Digital Control Module DDM De-orbit Deployment Mechanism DHS Data Handling System DHU Data Handling Unit DMM Design Maturity Margin DOA Degree of Adequacy of the Cost model DoD Depth of Discharge DoF Degree(s) of Freedom DSP Digital Signal Processor EADS European Aeronautic Defence and Space Company ECSS European Cooperation for Space Standardization ECU Electronic Control Unit EDAC Error Detection And Correction EDS Electronic Data Sheet EEE Electronic, Electrical and Electromechanical EEPROM Electrical Erasable PROM EESS Earth Exploration Satellite Services EGSE Electronic/Electrical Ground Support Equipment EM Engineering Model

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Acronym Definition EMC Electro Magnetic Compatibility EMI Electro Magnetic Interference EO Earth Observation EOL End Of Life EPE External Project Events EPROM Erasable PROM EQM Engineering Qualification Model ERC Earth Re-entry Capsule ESA European Space Agency ESOC European Space Operations Centre ESTRACK ESA’s ground station network EU European Union EVM Earned Value Management FAR Frame Analysis Report FCL Foldback Current Limiter FDIR Failure Detection Isolation and Recovery FDV Fill and Drain Valve FM Flight Model FMECA Failure Mode, Effects and Criticality Analysis FoV Field of View FPGA Field Programmable Gate Array FTA Fault Tree Analysis FVV Fill and Vent Valve G/S Ground Station GaAs Gallium Arsenide GaN Gallium Nitride GDIR General Design and Interface Requirements GEO Geostationary Orbit GNC Guidance, Navigations and Control GND Ground GNSS Global Navigation Satellite System

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Acronym Definition GPIO General Purpose Input / Output GPS Global Positioning System GS Ground Station GSE Ground Support Equipment GSP General Studies Programme GTO Geostationary Transfer Orbit HDDM Hold Down and Deployment Mechanism HDRM Hold-Down and Release Mechanism HGA High Gain Antenna HICDS Highly Integrated Computer & Data System HK Housekeeping HP Heat Pipe HPA High Power Amplifier HPC High Power Command HPTM High Priority TeleMetry HW Hardware I/O Input/Output ICD Interface Control Document ICDU Integrated Control and Data handling Unit IEEE Institute of Electrical and Electronics Engineers IFP Internal Final Presentation IO Input/Output IOD In-Orbit Demonstrator IQM Inherent Quality of the cost Model International Traffic in Arms Regulation (US law on export restrictions of defence and ITAR space related technology) ITU International Telecommunications Unit LCL Latching Current Limiter LEO Low Earth Orbit LEOP Launch and Early Operations Phase LGA Low Gain Antenna LNA Low Noise Amplifier

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Acronym Definition LSB Least Significant Bit LTAN Local Time of the Ascending Node LTCC Low Temperature Co-fired Ceramics MAP Multiplexed Access Point MEMS Micro Electro-Mechanical Systems MIPS Millions of Instructions Per Second MiSER Miniature Satellite Energy Regulating Radiator MLC Multi-Level Cell MLI Multi Layer Insulation MM Mass Memory MMM Mass Memory Module MMU Mass Memory Unit ModCod Modulation and Coding scheme MOI Moment of Intertia MOP Maximum Operating Pressure MPPT Maximum Peak Power Tracking MPT Mobile Phone based Transponder MPU Micro Processor Unit MSB Most Significant Bit NASA National Aeronautics and Space Administration NC Navigation Computer NCMM Navigation Computer and Mass Memory NRZ Non Return to Zero NTDM Nano-Terminator De-orbit Mechanism NVRAM Non-Volatile Random Access Memory OBC On Board Computer OBDH On Board Data Handling OBMU On Board Management Unit OBT On Board Time OCD Output Command Driver OTP One Time Programmable

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Acronym Definition OTS Off The Shelf P/L Payload PA Product Assurance PC Power Converter PCB Printed Circuit Board PCDU Power Conditioning and Distribution Unit PCM Power Conditioning Module PFM Proto Flight Model PM Processor Module PMD Propellant Management Device POE Project Owned Events PPS Pulse Per Second PROM Programmable ROM PSI Power Supply and I/O PSK Phase Shift Keying PTD Packet Telecommand Decoder PWM Pulse Width Modulation QIV Quality of the Input Values QoS Quality of Service QPSK Quadrature Phase Shift Keying R Reliability RAAN Right Ascension of the Ascending Node RAM Random Access Memory RCC Reconfiguration Command RE Reconfiguration Electronics RET Rover Elapsed Time RF Radio Frequency RFDU Radio Frequency Distribution Unit RM Reconfiguration Module ROM Read Only Memory RPE Relative Pointing Error

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Acronym Definition RS Reed Solomon RSA Relay Status Acquisition RT Remote Terminal RT Real-Time RTC Real-Time Clock RTD Resistance Temperature Detector RTU Remote terminal Unit RU Reconfiguration Unit S/C Space Craft SA Solar Array SADM Solar Array Drive Mechanism SBC Single Board Computer SBDL Standard Balanced Digital Link SBTA S-Band Transceiver Assembly SCCC Serially Concatenated Convolutional Codes SCET Spacecraft Elapsed Time SDM Solar Array Deployment Mechanism SDRAM Synchronous Dynamic Random Access Memory SE Saab Space AB SEE Single Event Effect SEL Single Event Latchup SET Single Event Transient SEU Single Event Upset SFT System Functional Test SGM Safe Guard Memory SIF Service Interface SM Structural Model SMA Shape Memory Alloy SME Small to Medium Enterprise SMU Spacecraft Management Unit SNR Signal to Noise Ratio

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Acronym Definition SoC System on a Chip SPF Single Point Failure SPT Scientific Payload Test SpW SpaceWire SR Space Research SRAM Static Random Access Memory SRRC Square Root Raised Cosine SS Sun Sensor SSMM Solid State Mass Memory SSO Sun-Synchronous Orbit SSPA Solid State Power Amplifier STM Structural Thermal Model SVF Software Verification Facility SW Software TB Transducer Bus TBC To be confirmed TBD To be determined TC TeleCommand TCR Temperature Coefficient of Resistance TM Telemetry TMC Telemetry and Telecommand TRL Technology Readiness Level TRP Technology Research Programme TSM Thermal Sensor Measurement TT&C Meaning of the acronym TTR Telemetry Telecommand and Reconfiguration board TTRS Telecommand, Telemetry, Reconfiguration and Safeguard memory module UART Universal Asynchronous Receiver/Transmitter UP/DW UP/DoWn US United States UT Unit Tester

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Acronym Definition UUT Unit Under Test VC Virtual Channel VCA Virtual Channel Assembler VCM Variable Coding and Modulation WD Watch Dog XBTA X-Band Transmitter Assembly ∆V Delta V: Change in Velocity