contributions to the advance of the integration density of cubesats · instute of aeronaucs and...

Institute of Aeronautics and Astronautics: Scientific Series Band 3

Universitätsverlag der TU Berlin

Sebastian Grau

Contributions to the Advance of the Integration Density of CubeSats

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Sebastian Grau

Contributions to the Advance of theIntegration Density of CubeSats

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The scientific serie Institute of Aeronautics and Astronautics:Scientific Series of the Technische Universität Berlin is edited by:Prof. Dr.-Ing. Dieter Peitsch,Prof. Dr.-Ing. Andreas Bardenhagen,Prof. Dr.-Ing. Klaus Brieß,Prof. Dr.-Ing. Robert Luckner,Prof. Dr.-Ing. Julien Weiss

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Institute of Aeronautics and Astronautics: Scientific Series | 3

Sebastian Grau

Contributions to the Advance of theIntegration Density of CubeSats


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Bibliographic information published by the Deutsche NationalbibliothekThe Deutsche Nationalbibliothek lists this publication in theDeutschen Nationalbibliografie; detailed bibliographic data areavailable on the Internet at http://dnb.dnb.de.

Universitätsverlag der TU Berlin, 2019http://verlag.tu-berlin.de

Fasanenstr. 88, 10623 BerlinTel.: +49 (0)30 314 76131 / Fax: -76133E-Mail: [email protected]

Zugl.: Berlin, Techn. Univ., Diss., 2018Gutachter: Prof. Dr.-Ing. Klaus BrießGutachter: Prof. Dr.-Ing. Hakan Kayal (Universität Würzburg)Die Arbeit wurde am 30. August 2018 an der Fakultät V unterVorsitz von Prof. Dr.-Ing. Andreas Bardenhagen erfolgreich verteidigt.

This work – except for quotes, figures and where otherwise noted –is licensed under the Creative Commons Licence CC BY 4.0http://creativecommons.org/licenses/by/4.0/

Cover image:Ausschnitt: NASAhttps://images-assets.nasa.gov/image/iss047e120450/iss047e120450~orig.jpgPublic domain

Print: Pro BUSINESSLayout/Typesetting: Sebastian Grau, Karsten Gordon

ISBN 978-3-7983-3026-9 (print)ISBN 978-3-7983-3027-6 (online)

ISSN 2512-5141 (print)ISSN 2512-515X (online)

Published online on the institutional Repositoryof the Technische Universität BerlinDOI 10.14279/depositonce-7293http://dx.doi.org/10.14279/depositonce-7293

https://images-assets.nasa.gov/image/iss047e120450/iss047e120450~orig.jpg

http://dx.doi.org/10.14279/depositonce-7293

http://creativecommons.org/licenses/by/4.0/



Mailto: [email protected]

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Contents

1 Introduction 11.1 Recent and Future Evolution of CubeSat Launches . . . . . 2

1.1.1 NewSpace Constellations . . . . . . . . . . . . . . . 21.1.2 Civil Constellations . . . . . . . . . . . . . . . . . . 31.1.3 University-Class Spacecraft . . . . . . . . . . . . . . 4

1.2 Thesis Outline . . . . . . . . . . . . . . . . . . . . . . . . . 5

2 High Performance Single Unit CubeSat Design Approaches 92.1 CubeSat Market Observations . . . . . . . . . . . . . . . . . 9

2.1.1 Independent Commercial Spacecraft . . . . . . . . . 102.1.2 Commercially-Procured Spacecraft . . . . . . . . . . 112.1.3 Independent University Spacecraft . . . . . . . . . . 17

2.2 High Performance Single Unit CubeSat Design Criteria . . . 232.2.1 Provided Payload Resources . . . . . . . . . . . . . 232.2.2 Attitude Determination and Control Capabilities . . . 232.2.3 Downlink Capabilities . . . . . . . . . . . . . . . . . 242.2.4 Data and Power Bus . . . . . . . . . . . . . . . . . 24

2.3 Highly Integrated Solar Panel Design Criteria . . . . . . . . 25

3 Picosatellite Solar Antenna 273.1 Solar Antenna State of the Art . . . . . . . . . . . . . . . . 273.2 Patch Antenna Development . . . . . . . . . . . . . . . . . 293.3 Solar Cell Integration . . . . . . . . . . . . . . . . . . . . . 333.4 Solar Antenna Functional Verification . . . . . . . . . . . . 35

4 Magnetic Actuator Optimization 374.1 Magnetic Actuator State of the Art . . . . . . . . . . . . . . 38

4.1.1 Wound Torque Rods . . . . . . . . . . . . . . . . . 394.1.2 Embedded Air Coils . . . . . . . . . . . . . . . . . . 394.1.3 Wound Air Coils . . . . . . . . . . . . . . . . . . . . 41

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vi Contents

4.2 Magnetorquer Optimization State of the Art . . . . . . . . . 424.2.1 Experiment-Based Optimization . . . . . . . . . . . 434.2.2 Model-Driven Optimization . . . . . . . . . . . . . . 434.2.3 Sequential Quadratic Programming . . . . . . . . . . 44

4.3 Formulation of a Novel Optimization Procedure . . . . . . . 454.3.1 Mathematical Modeling of Magnetic Coil Properties . 464.3.2 High-Dimensional Parameter Space Implementation . 474.3.3 Discrete and Integer Input Parameters . . . . . . . . 484.3.4 Flexible Optimization Objectives Definition . . . . . 484.3.5 Global Optimum Search . . . . . . . . . . . . . . . . 50

4.4 Optimization Results . . . . . . . . . . . . . . . . . . . . . 524.4.1 Wound Torque Rod Optimization Results . . . . . . 524.4.2 Wound Air Coil Optimization Results . . . . . . . . 524.4.3 Embedded Air Coil Optimization Results . . . . . . . 54

5 A New Class of Picosatellite Attitude Actuators 555.1 Technological Evolution of Fluid Spacecraft Actuators . . . . 555.2 Objectives for Actuator Miniaturization . . . . . . . . . . . . 575.3 Fluid-Dynamic Actuator Fundamentals . . . . . . . . . . . . 59

5.3.1 Planar Actuators . . . . . . . . . . . . . . . . . . . 605.3.2 Three-Dimensional Actuators . . . . . . . . . . . . . 635.3.3 General Spacecraft Dynamics . . . . . . . . . . . . . 67

5.4 Fluid Actuator Conduits for CubeSat Applications . . . . . . 685.4.1 Conduit Considerations . . . . . . . . . . . . . . . . 685.4.2 Pump Housing . . . . . . . . . . . . . . . . . . . . . 695.4.3 Actuator Electronics . . . . . . . . . . . . . . . . . . 705.4.4 Manufacturing of Monolithic, Integrated Conduits . . 705.4.5 First Rapid Prototyping Experiences . . . . . . . . . 715.4.6 Conduit Geometries for CubeSat Applications . . . . 725.4.7 Advantages of Conduit Rapid Prototyping . . . . . . 74

5.5 Driver Electronics . . . . . . . . . . . . . . . . . . . . . . . 745.5.1 Electronics Miniaturization . . . . . . . . . . . . . . 745.5.2 Flexible Development Platform . . . . . . . . . . . . 76

5.6 Functional Verification . . . . . . . . . . . . . . . . . . . . 785.6.1 Power Consumption . . . . . . . . . . . . . . . . . . 805.6.2 Dynamical Properties . . . . . . . . . . . . . . . . . 80

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Contents vii

5.7 Fluid-Dynamic Actuators and Reaction Wheels . . . . . . . . 855.7.1 Angular Rate and Acceleration . . . . . . . . . . . . 865.7.2 Time-Optimal Slew Maneuvers . . . . . . . . . . . . 865.7.3 Analysis of Traveled Angles . . . . . . . . . . . . . . 87

5.8 Redundancy Concepts . . . . . . . . . . . . . . . . . . . . . 895.8.1 L-Shaped Conduits . . . . . . . . . . . . . . . . . . 895.8.2 Crown-Shaped Conduits . . . . . . . . . . . . . . . . 92

6 A Highly Integrated Single Unit CubeSat Solar Panel 956.1 Current Integration Density of CubeSat Solar Panels . . . . 98

6.1.1 Mechanical Properties . . . . . . . . . . . . . . . . . 996.1.2 Power Generation . . . . . . . . . . . . . . . . . . . 1006.1.3 Attitude Determination Sensors . . . . . . . . . . . . 1016.1.4 Attitude Control Actuators . . . . . . . . . . . . . . 1026.1.5 Harness . . . . . . . . . . . . . . . . . . . . . . . . 1036.1.6 Conclusion . . . . . . . . . . . . . . . . . . . . . . . 104

6.2 Design of a Highly Integrated, Multi-Functional Solar Panel . 1056.2.1 Communication . . . . . . . . . . . . . . . . . . . . 1066.2.2 Power Generation and Distribution . . . . . . . . . . 1066.2.3 Attitude Determination . . . . . . . . . . . . . . . . 1086.2.4 Attitude Actuators . . . . . . . . . . . . . . . . . . . 1086.2.5 Command and Data Handling . . . . . . . . . . . . 110

6.3 Solar Panel Assembly and Test . . . . . . . . . . . . . . . . 1106.3.1 Assembly . . . . . . . . . . . . . . . . . . . . . . . 1116.3.2 Functional Verification . . . . . . . . . . . . . . . . 1136.3.3 Environmental Tests . . . . . . . . . . . . . . . . . . 114

6.4 Highly Integrated Multi-Functional Solar Panel Advantages . 116

7 Summary and Conclusion 1197.1 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1197.2 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . 121

A Magnetic Coil Data 137

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List of Figures

1.1 Number of CubeSat launches 2003–2022 . . . . . . . . . . . 11.2 Distribution of civil CubeSat launches 2003–2022 . . . . . . 31.3 Distribution of university CubeSat application over form factor

2003–2018 . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

3.1 Radiation mechanism of a patch antenna and integration of asolar cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

3.2 Return loss and axial ratio of a quasi-square patch antenna fordifferent thicknesses of the dielectric medium . . . . . . . . 30

3.3 Return loss and axial ratio of all researched patch antennas . 313.4 Right hand and left hand circular polarization antenna gains

for all researched patch geometries in the xz plane in dB . . 323.5 Right hand and left hand circular polarization antenna gains

for all researched patch geometries in the yz plane in dB . . 323.6 Cross section schematic of solar antenna . . . . . . . . . . . 333.7 Image of the fully integrated solar patch antenna . . . . . . 343.8 Solar antenna return loss measurement . . . . . . . . . . . . 353.9 Solar antenna gain measurement . . . . . . . . . . . . . . . 36

4.1 Overview of magnetic actuator properties . . . . . . . . . . 404.2 Overview of optimized magnetic actuator properties . . . . . 53

5.1 First demonstrator of a CubeSat fluid-dynamic actuator on anair bearing test facility . . . . . . . . . . . . . . . . . . . . . 58

5.2 Circular conduit . . . . . . . . . . . . . . . . . . . . . . . . 615.3 Rectangular conduit . . . . . . . . . . . . . . . . . . . . . . 635.4 L-shaped conduit . . . . . . . . . . . . . . . . . . . . . . . 645.5 Crown-shaped conduit . . . . . . . . . . . . . . . . . . . . . 665.6 First monolithic, integrated fluid-dynamic actuator conduit . 725.7 Monolithic, integrated fluid-dynamic actuator conduit . . . . 73

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x List of Figures

5.8 Circular integrated conduit used for student projects . . . . . 735.9 Block diagram of initial pFDA electronics . . . . . . . . . . 755.10 First fully-integrated picosatellite fluid-dynamic actuator opti-

mized for CubeSat application . . . . . . . . . . . . . . . . 765.11 Components of the flexible development platform . . . . . . 775.12 Flexible development platform assembled for air bearing use 775.13 First demonstrator of a pFDA using a 3D-printed conduit . . 795.14 Angular rate measurements of a pFDA . . . . . . . . . . . . 795.15 Pump power demand at 5 V in mW . . . . . . . . . . . . . . 815.16 Pump power demand at 12 V in mW . . . . . . . . . . . . . 815.17 Actuator parameter estimation based on a first-order LTI model 835.18 Estimated actuator parameters . . . . . . . . . . . . . . . . 845.19 Comparison of simulated actuator properties . . . . . . . . . 875.20 Comparison of simulated time-optimal slew maneuvers . . . 885.21 Maximum angles traveled using time-optimal and non-optimal

maneuvers . . . . . . . . . . . . . . . . . . . . . . . . . . . 895.22 L-shaped conduits in a redundant configuration . . . . . . . 905.23 Experimental assembly of redundant fluid-dynamic actuator

configuration based on L-shaped conduits . . . . . . . . . . 915.24 Crown-shaped conduits in a redundant configuration . . . . 93

6.1 Schematic view of a proposed high density single unit CubeSatusing highly integrated, multi-functional solar panels . . . . 96

6.2 Nodes of a proposed high density single unit CubeSat . . . . 976.3 Solar cell orientation and mounting hole placement . . . . . 1006.4 Allocation of antenna and solar cell pads . . . . . . . . . . . 1076.5 Magnetic coil layout . . . . . . . . . . . . . . . . . . . . . . 1096.6 Pre-assembled multi-functional solar panel . . . . . . . . . . 1116.7 Fully assembled multi-functional solar panel . . . . . . . . . 1126.8 Fit-check of the fully assembled multi-functional solar panel 1136.9 Solar panels mounted for mechanical tests . . . . . . . . . . 1156.10 Damaged solar panel and torn-off solar antenna . . . . . . . 1156.11 CAD model of a high performance single unit CubeSat . . . 118

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List of Tables

2.1 Integrated CubeSat platforms: mechanical and electrical char-acteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2.2 Integrated CubeSat platforms: attitude determination andcontrol system characteristics . . . . . . . . . . . . . . . . . 16

2.3 Integrated CubeSat platforms: communication system charac-teristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

5.1 Comparison of dynamical properties of CubeSat reaction wheelsand the picosatellite fluid-dynamic actuator . . . . . . . . . 86

6.1 Single unit CubeSat solar panel properties . . . . . . . . . . 996.2 Single unit CubeSat solar panel components . . . . . . . . . 1026.3 Components of a proposed high-density single unit CubeSat

design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116

A.1 Wound torque rod data . . . . . . . . . . . . . . . . . . . . 137A.2 Embedded air coil data . . . . . . . . . . . . . . . . . . . . 138A.3 Wound air coil data . . . . . . . . . . . . . . . . . . . . . . 139

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Glossary

AraMiS

Italian for Modular Architecture for Satellites ; a project to researchsmall satellite modularization, conducted at Politecnico Torino, Italy.103, 105

Astrofein

Astro- und Feinwerktechnik Adlershof GmbH, a Berlin, Germany basedmanufacturer of satellite components. 86

BEESAT

Berlin Experimental and Educational SATellite (BEESAT), a series ofCubeSats designed and operated by Technische Universität Berlin. Alsothe name of the first single unit CubeSat mission at TU Berlin, whichwas designed to verify the RW 1 reaction wheel in space. 17–19, 38,41, 54, 72, 100, 103, 104, 106, 108, 112

BEESAT-2

The second in the series of BEESAT satellites, designed to demonstrateadvanced attitude determination and control capabilities. 17, 18

BEESAT-4

The fourth in the series of BEESAT satellites, designed to demonstrateorbit determination using GPS and act as a target satellite for BIROS.6, 17–19, 24, 25, 117, 118

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xiv Glossary

BEESAT-5 to BEESAT-8

A swarm of four quarter unit CubeSats in the series of BEESAT satellites,designed to demonstrate novel UHF transceivers and GPS receivers;planned to be launched together from a single unit deployment containerin 2019. 5, 6, 17, 21, 22, 24, 25, 98

BIROS

The Bi-spectral Infrared Optical System (BIROS) is successor to theTET-1 satellite and part of the DLR FireBIRD mission. It releasedBEESAT-4 into orbit from an internal deployment container. 17

BK77

UHF transceiver used on the BEESAT to BEESAT-4 missions. 17, 18

Blue Canyon Technologies

Blue Canyon Technologies Inc. is a Boulder, Colorado based CubeSatvendor. 11, 15

CanX-1

Canadian Advanced Nanospace eXperiment 1 (CanX-1) was launchedas one of the first CubeSats ever in 2003. 41

Clyde Space

Clyde Space is a Scottish microsatellite and CubeSat supplier. 11, 14,15

COMPASS-1

Second German CubeSat in space, developed at FH Aachen, Germany.42

CubePMT

The CubeSat Power Management Tile is a multi-functional highly in-tegrated CubeSat solar panel developed at Politecnico Torino, Italy, inthe scope of the AraMiS-C1 project. 105

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Glossary xv

Dove

The Dove spacecraft are triple unit CubeSats, built and operated byAmerican Earth observation company Planet Labs in the Flock constel-lations. 2, 10–12, 22, 95

ECSS

European Cooperation for Space Standardization. 114

Flock

Earth observation satellite constellations consisting of Dove satellites,developed and operated by Planet Labs. 10

GaAs

Gallium arsenide, a compound material used for the production of highefficiency solar cells. 28

GNU Octave

Free software that features a high-level programming language, primarilyintended for numerical computations. 47, 48, 50, 75, 108

GomSpace

GomSpace is a Danish CubeSat and nanosatellite manufacturer basedin Aalborg, Denmark. 11, 14

GOMX-3

In-orbit demonstration mission developed by GomSpace and funded byESA. 11, 20

HISPICO

Highly Integrated S Band Transmitter for Pico and Nano Satellites devel-oped by Berlin based company iQ spacecom and Technische UniversitätBerlin. 24, 29

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xvi Glossary

ISIS

Innovative Solutions In Space B.V. (ISIS) is a Dutch nanosatellitecompany located in Delft, The Netherlands. 11, 14

Kepler Communications

Kepler Communications Inc. is a Canadian satellite communicationcompany located in Toronto, Canada. 2, 11

KiCAD

KiCAD is a free and open source electronic design automation software.108

Ku-band

Frequency band in the range from 10.7–17.5 GHz. 2

Lemur-2

The Lemur-2 satellites are triple unit CubeSats, built and operated byAmerican company Spire Global. 2, 11

MicroMAS

Micro-sized Microwave Atmospheric Satellite (MicroMAS) is a three-unitCubeSat developed by Massachusetts Institute of Technology SpaceSystems Laboratory. 4

MicroMAS 2

The Micro-sized Microwave Atmospheric Satellite 2 (MicroMAS 2)mission consists of two three-unit CubeSats developed by MassachusettsInstitute of Technology Space Systems Laboratory and is the successorto the MicroMAS spacecraft. 4

MOVE-II

Munich Orbital Verification Experiment II (MOVE-II) is a single unitCubeSat developed at TU München and launched in 2018. 41

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Glossary xvii

MSSL

Mullard Space Science Laboratory. 4

NASA

National Aeronautics and Space Administration. 4, 11, 12, 56

OUFTI 1

Orbital Utility For Telecommunications/Technology Innovations 1 (OUFTI1) is a single unit CubeSat built by students of the Université de Liège,Belgium. 41

PC/104

Embedded computer standard defining both form factors and computerbuses intended for extreme environment applications. 12, 15, 18–20,23, 24, 95

PEEK

Polyether Ether Ketone, a high-temperature, high-strength thermoplasticpolymer. 71

PID

Control law that uses a proportional, integral, and derivative componentto calculate the output. 56

PiNaSys

TU Berlin project: Development and Experimental Testing of Miniatur-ized Components for Distributed Pico- and Nanosatellite Systems. 21,23

PiNaSys II

TU Berlin project: Further Development and Verification of MiniaturizedComponents for Distributed Pico- and Nanosatellite Systems (BEESAT-5to BEESAT-8). 5, 21, 23, 96

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xviii Glossary

Planet Labs

American Earth observation company, developing and operating theFlock constellations formed by a large number of Dove satellites. 2,10–12, 22, 95

Pumpkin

Pumpkin Space Systems, American CubeSat component supplier basedin San Francisco, California. 11

PVC

Polyvinyl Chloride, plastic polymer used among others for electric cableisolation. 57

QB50

QB50 is a multi-satellite project to study the lower and middle thermo-sphere with a swarm of up to 50 double and triple unit CubeSats. 4,119

RAVAN

Radiometer Assessment using Vertically Aligned Nanotubes (RAVAN) isa triple unit CubeSat developed at Johns Hopkins University, Baltimore,Maryland, based on the XB3 CubeSat bus of Blue Canyon Technologies.11, 12

REXUS

Rocket Experiments for University Students. 92

REXUS/BEXUS

Rocket/Balloon Experiments for University Students. 92

RW 1

Picosatellite reaction wheel offered by Astro- und Feinwerktechnik Adler-shof GmbH, Berlin, Germany. 17, 20, 86–90

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Glossary xix

S-band

Frequency band in the range from 2–4 GHz. 5, 6, 10, 15, 22, 24, 26–29,120

SI

The International System of Units. 48

Sky and Space

Sky and Space Global Ltd, UK-based satellite telecommunication com-pany. 2, 11

SMA

SubMiniature version A, a type of coaxial RF connectors. 33, 106

S-Net

S-Band Netzwerk für kooperierende Satelliten, German for S-bandnetwork of cooperating satellites, a non-CubeSat standard nanosatelliteformation developed and operated by Technische Universität Berlin. 4,17

Spire Global

Spire Global, Inc is a San Francisco, California based NewSpace company.2, 11

SSTL

Surrey Satellite Technology Ltd. (SSTL), British small satellite manu-facturer based in Guildford, United Kigdom. 11

TechnoSat

A nanosatellite mission developed at Technische Universität Berlin fordemonstration of the TUBiX20 satellite bus. 17, 57

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xx Glossary

TROPICS

NASA’s Time-Resolved Observations of Precipitation structure andstorm Intensity with a Constellation of Smallsats mission. 4

TU Berlin

Technische Universität Berlin. 4–6, 17, 18, 21, 24, 29, 38, 41, 57, 74,96, 97, 108, 110, 113, 114, 120, 121

TU Dresden

Technische Universität Dresden. 4

TU München

Technische Universität München. 41

TUBIN

A TU Berlin remote fire detection mission based on the TUBiX20satellite bus, successor to the TechnoSat mission. 17

TUBiX20

20 kg nanosatellite platform developed at TU Berlin, utilized for themissions TechnoSat and TUBIN. 24, 95–98

TUPEX-6

TU Berlin Picosatellite Experiment 6, a free-falling unit deployed from aREXUS sounding rocket to demonstrate redundancy concepts featuringfluid-dynamic actuators for picosatellites. 92

Tyvak

Tyvak Nano-Satellite Systems, Inc. is an American nanosatellite manu-facturer located in Irvine, California. 14

UHF

Ultra high frequency band. 5, 10, 15–18, 20, 22, 24, 27, 117

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Glossary xxi

UiO

University of Oslo. 4

USART

Universal Synchronous/Asynchronous Receiver Transmitter. 75

UWE-3

Universität Würzburg Experimentalsatellit 3. 6, 41, 104

VHF

Very high frequency band. 16, 27

XB3

Single unit integrated CubeSat platform for use in triple unit CubeSatsoffered by Blue Canyon Technologies Inc. 11, 13

X-band

Frequency band in the range from 8–12 GHz. 10, 15, 22

ZARM Technik AG

Bremen, Germany based manufacturer of magnetorquers and magne-tometers. 39

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Acronyms

ADCS Attitude Determination and Control System. 20, 22, 36–38, 59,91, 101, 103, 104, 108

CAD Computer Aided Drawing. 85, 117

CAN Controller Area Network. 18, 21, 22, 24–26, 98, 110

CDS CubeSat Design Specification. 9, 12

CMOS Complementary Metal-Oxide-Semiconductor. 101

COTS Commercial Off-The-Shelf. 9, 18, 37

DC Direct Current. 28, 70, 77, 106

DFG German Research Foundation, from German Deutsche Forschungs-gemeinschaft. 6

DLR German Aerospace Center, from German Deutsches Zentrum fürLuft- und Raumfahrt . 17, 92

EAC Embedded Air Coil. 37–39, 41, 43–45, 48, 49, 54, 100, 102, 114,120, 121

EDA Electronic Design Automation. 108

EEI Earth Energy Imbalance. 11

ESA European Space Agency. 92

FAB Front Access Board. 20, 21

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xxiv Acronyms

FDA Fluid-Dynamic Actuator. 55, 57–59, 121

FFC Flexible Flat Cable. 103, 104

FFU Free-Falling Unit. 92, 93

FIPEX Flux-Φ-Probe EXperiment. 4

GPIO General Purpose Input and Output. 20

GPS Global Positioning System. 5, 10, 17–19, 22, 28, 93

HPBW Half Power Beamwidth. 31, 33, 36

I2C Inter-Integrated Circuit. 20, 21, 75, 110

IDC Insulation-Displacement Contact. 18, 25, 103, 104

IM Instant Messaging. 2

IMU Inertial Measurement Unit. 10

INMS Ion-Neutral Mass Spectrometer. 4

IOC In-Orbit Checkout. 3

IOD In-Orbit Demonstration. 17, 22, 105

IoT Internet of Things. 2, 3

ISL Inter-Satellite Link. 2, 5, 19

IZM Institute for Reliability and Microintegration, from German Institutfür Zuverlässigkeit und Mikrointegration. 6, 27, 35, 110, 111, 113,114, 120

LEO Low Earth Orbit. 56

LHCP Left Hand Circular Polarization. 31

LTI Linear Time-Invariant. 81, 82

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M2M Machine-To-Machine. 2, 3

MEMS Micro-electro-mechanical System. 10, 17, 24, 70, 101, 108

MKI Multi-functional Integration of Miniaturized Satellite Componentsfor Increased Payload Capacity of Pico-Satellites. 6, 27

m-NLP Multi-Needle Langmuir Probe. 4

MPPT Maximum Power Point Tracker. 25, 34, 101, 104, 105, 107, 110,113

OBC On-Board Computer. 17, 18, 22, 97, 98, 104, 108, 116, 117

OOV On-Orbit Verification. 28

PCB Printed Circuit Board. 12, 21, 22, 25, 35–38, 43, 50, 54, 60, 70,75, 77, 93, 94, 96, 97, 99, 102, 105, 108, 110, 111

PCDU Power Control and Distribution Unit. 17, 18, 22, 97, 98, 101, 103,104, 107, 110, 116, 117

PCU Power Control Unit. 18, 22

PDH Payload Data Handling. 17, 18, 25, 26

pFDA Picosatellite Fluid-Dynamic Actuator. 7, 59, 62, 67, 68, 72, 74–78,81, 86–92, 97, 98, 105, 107, 108, 110–112, 117, 118, 120–122

PSD Position-Sensitive Device. 101, 104, 108

PTT Push-To-Talk. 2

RBF Remove Before Flight. 20, 100, 101

RF Radio Frequency. 28, 30, 33, 35, 93, 106

RHCP Right Hand Circular Polarization. 31, 36

RWA Reaction Wheel Assembly. 18, 20

SLS Selective Laser Sintering. 70, 71

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xxvi Acronyms

SNSB Swedish National Space Board. 92

SQP Sequential Quadratic Programming. 42, 44, 45, 47

UART Universal Asynchronous Receiver Transmitter. 20, 21

UK United Kingdom. 2

UWE-3 University of Würzburg Experimental satellite, from German Uni-versität Würzburg Experimentalsatellit 3 . 20, 21, 24, 96

WAC Wound Air Coil. 37, 38, 41, 43, 52, 103

WDE Wheel Drive Electronic. 18, 20

WTR Wound Torque Rod. 37–39, 41, 43, 44, 52

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Latin Symbols

𝐴 area in m2. 46, 50, 61, 62, 65

𝑎 board length of an embedded air coil in m. 48

𝐵 magnetic field in N/(m A). 114

𝑑c core diameter in m. 46

𝑑w wire diameter in m. 46

𝐻 Angular momentum in kg m2 s. 61–68, 84, 85, 90, 93

𝐻 Angular momentum vector in kg m2 s. 59, 60, 65–68

ℎ height in m. 99

𝐼 current in A. 46

𝐼 scalar moment of inertia in kg m2. 19, 62, 84–86

𝐼 moment of inertia tensor in kg m2. 60

𝐼s inertia of a spacecraft in kg m2. 67

𝑙 length in m. 19, 99

𝑙c core length in m. 46

𝑙w wire length in m. 46

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xxviii Latin Symbols

𝑚 number of layers of a wound air coil. 50

𝑀 mass in kg. 19, 39, 46, 49, 59–62, 99, 137–139

𝑝 momentum vector in kg m/s. 59

𝑃 electric power in W. 39, 46, 49, 99, 137–139

𝑟 radius vector in m. 61, 62

𝑟 radius vector in m. 59–61

𝑅 electric resistance in W. 46

𝑆 cross-section in m2. 61–66, 90, 93

𝑇 time constant of a first-order, linear time-invariant system in s. 82

𝑡 time in s. 82

𝑇 torque distribution matrix. 68

𝑈 voltage in V. 46, 48, 137–139

𝑣 velocity in m/s. 61–66, 90, 93

𝑣 velocity vector in m/s. 59–61

𝑤 length in m. 99

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Greek Symbols

𝛼 specific resistance temperature Coefficient. 46

𝜇 magnetic dipole in A m2. 39, 46, 49, 114, 137–139

𝜇0 vacuum permeability. 114

𝜌 density in kg/m3. 46, 61–66, 90, 93

𝜏 torque vector in N m. 67

𝜏 torque in N m. 83–85

𝜏 𝑑 disturbance torque vector in N m. 67

𝜔 angular rate vector °/s. 60

𝜔 angular rate °/s. 62, 82–86

𝜔𝑠 angular rate vector of a satellite in °/s. 67

�̇� angular acceleration in °/s2. 82–84, 86

�̇�𝑠 angular acceleration vector of a satellite in °/s. 67

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1 Introduction

In less than fifteen years, CubeSats developed from a tool in space engineeringeducation to a platform able to support low-cost scientific and Earth observa-tion missions [3, 4]. They became key to a novel class of commercial missionconcepts using large numbers of CubeSats in constellations to gather datawith high temporal and spatial resolution [5, 6]. This lead to a sharp increasein the number of CubeSat launches over the past years, which is predicted tocontinue (cf. [1, 2], figure 1.1).ii

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Figure 1.1: Number of CubeSat launches 2003–2022 (based on [1, 2])

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2 1 Introduction

1.1 Recent and Future Evolution of CubeSat Launches

Williams, Doncaster, and Shulman found that "overall sizes in the nanosatellitemarket are increasing to accommodate demand for additional payload capabil-ity", and that "the 3 U form factor is still expected to remain the standard inthe market over the next five years" [2].

1.1.1 NewSpace Constellations

Williams et al. identify U.S. Earth observation company Planet Labs to be thesingle biggest contributor to nanosatellite launches [2]. The Dove satellitesare triple unit CubeSats that provide complete coverage of the Earth withan optical resolution of 3–5 m and daily revisits [6]. Planet Labs alone isresponsible for 35 % of all nanosatellite launches between 2008 and 2017 [2].

With Spire Global exists a second company that operates an establishedconstellation of triple unit CubeSats [7]. Until February 2018, Spire Globalsuccessfully launched 67 of their Lemur-2 satellites. The forecast in [2] statesthe expectation, that Earth observation and remote sensing constellations willmake up 50 % of the nanosatellite market, and communications constellationswill account for 20 % over the next five years.

Among the new communication constellations is UK-based company Skyand Space who launched an inter-satellite link and communications proof-of-concept mission consisting of three triple unit CubeSats, called ThreeDiamonds, in 2017 [8]. The mission demonstrates "a number of functionalities,including phone calls, internet of things (IoT), machine-to-machine (M2M),instant messaging (IM), push-to-talk (PTT) services, and data store-and-forward between different locations on Earth", according to the company[9]. In 2017, Sky and Space announced the transition towards a mega-constellation of 200 6 U CubeSats, which is "expected to be fully deployedand operational by 2020" [10]. The so-called Pearl spacecrafts will form amobile telecommunication network for the equatorial countries.

In January 2018, Canadian company Kepler Communications launched theirfirst triple unit CubeSat as a demonstrator for their planned Ku-band commu-

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1.1 Recent and Future Evolution of CubeSat Launches 3

nications constellation [11]. The complete constellation will consist of 75–140operational spacecraft [12].

Williams et al. identify further companies, that are developing communicationsconstellations "centered around serving the rapidly growing IoT/M2M market"[2]. In-orbit checkout of constellations developed by companies like Astrocast[13], Aerial & Maritime [14], or SRT Marine Systems [15] is expected for 2018or 2019 [1, 2].

1.1.2 Civil Constellations

With regard to civil CubeSat missions, Williams et al. state that "despite rapidcommercial market share growth, civil1 (...) operator demand is expected toremain consistent over the next 5 years" [2]. This statement is contrasted bydata on civil launches taken from [1] and shown in figure 1.22: The announcedCubeSat launches for 2018 triple the number of launches in 2017.

1Williams et al. define civil operators as "operators whose primary satellite purpose isnon-military or non-profit activities" [2].

2The eight operator classes found in [1] have been regrouped according to the threeoperator classes found in [2].i

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Figure 1.2: Distribution of civil CubeSat launches 2003–2022 (based on [1])

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4 1 Introduction

Among those announced missions is NASA’s TROPICS mission; a scientificconstellation of six identical triple unit CubeSats, which is currently underdevelopment and "on-track to deliver fligh-ready hardware in 2019", as statedby Blackwell, Burianek, Clark, et al. in [16]. The spacecraft is based on theMicroMAS [17] and MicroMAS 2 [16] missions, launched in July 2014 andJanuary 2018, respectively. All satellites feature a commercially-procured bus,and a single unit payload for observation of precipitation structure and stormintensity of tropical cyclones [16].

The largest constellation of civil operated CubeSats is QB50 [5], consistingof a total of 36 double and triple unit spacecraft launched in two batches inApril and June of 2017 [18, 19]. In contrast to other CubeSat constellations,satellites were contributed by multiple universities and research organizations.Each spacecraft hosts only one out of the three common science units. Intotal, 13 ion-neutral mass spectrometers (INMSs) provided by MSSL, 19flux-Φ-probe experiment (FIPEX) sensors provided by TU Dresden, and 11multi-needle Langmuir probes (m-NLPs) supplied by UiO are located on thespacecraft of the constellation and will be used to study the properties of theupper thermosphere [20]. Each of the science units was developed to "half aCubeSat unit volume budget (excluding forward protuberance)" [21].

1.1.3 University-Class Spacecraft

Swartwout and Jayne, analyzing university-class spacecraft trends in [22],define three categories of university-class programs: flagship universities thatfly reliable and significant missions every few years, independent universitiesthat have developed their own string of successful missions, and hobbyistswith low flight rates and high on-orbit failure rates. In the scope of their work,TU Berlin is considered to be among the first flagship universities worldwide.By that definition, TU Berlin is constantly pushing the state of the art ofsmall satellite technology.

The first satellite constellation of TU Berlin, called S-Net, was successfullybrought into orbit on February 1, 2018 [23]. The four satellites have dimensionsof 25 × 25 × 25 cm [24] and are therefore not compliant with the CubeSat

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1.2 Thesis Outline 5

standard [25]. Primary mission objective is the demonstration of multi-point inter-satellite link in the S-band between the members of a swarm ofnanosatellites without active orbit control [24].

TU Berlin’s second constellation, the BEESAT-5 to BEESAT-8 satellitesof the PiNaSys II mission, are scheduled for launch in 2018. The foursatellites are quarter-unit CubeSats intended for demonstration of "a newlydeveloped communications subsystem in the UHF band and an experimentalGPS receiver", as stated by Baumann et al. in [26]. Beyond that, the satellitesare designed almost complete single-fault tolerant and feature multi-functionaloptical attitude determination sensors [27]. Another specialty of the missionis, that the four satellites will be launched as a pack from one single unitCubeSat deployer [28].

1.2 Thesis Outline

In conclusion, the economic, civil, and university sectors show a clear trendtowards more CubeSat missions to be launched and an increase in payloadperformance. University missions are starting to put more attention onscientific CubeSat missions (cf. [5], figure 1.3) At the same time, ongoingi

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Figure 1.3: Distribution of university CubeSat application over form factor 2003–2018 (based on [1])

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6 1 Introduction

research on miniaturization of CubeSat subsystems allows to develop smallerbuses. Combining these factors, the question arises to what extent it is possibleto miniaturize the bus of a CubeSat, developed in an university environment,in order to increase available resources for high performance payloads.

In this context, this thesis investigates a concept using highly integrated,multi-functional solar panels to increase the overall integration density ofCubeSats. Research presented in this work was carried out in the scope of ajoint, DFG-funded project titled Multi-functional Integration of MiniaturizedSatellite Components for Increased Payload Capacity of Pico-Satellites (MKI)at the Chair of Space Technology at TU Berlin and Fraunhofer Institute forReliability and Microintegration (IZM). Authorship of original work of theauthor or the project partners is highlighted in the following outline.

In chapter 2, the author discusses criteria and drivers for high performanceCubeSat design. Three different CubeSat categories are identified from marketobservations, and their design characteristics are compared against eachother. In this context, special attention is paid to missions of the two majorGerman contributors to CubeSat development, launch, and operation: TUBerlin and Universität Würzburg. Analysis of the BEESAT-4, BEESAT-5to BEESAT-8, and UWE-3 missions together with the results of the marketobservations are used to derive design criteria for high performance single unitCubesats. One approach to meet those criteria is the use of highly integrated,multi-functional solar panels, which host components from different satellitesubsystems. The author ends this chapter with the definition of design criteriafor multi-functional solar panels.

To show, how formulated design criteria are met, each of the following chaptersrepresents stand-alone research and development of subsystem componentsthat reside on the multi-functional solar panel. Each chapter begins witha discussion of the state of the art in the relevant field. Subsequently,development process, implementation, and results are described.

Chapter 3 summarizes research on solar antennas for S-band communicationfrom space to the ground. Work was carried out at Fraunhofer IZM, andsupported by the author in an advisory capacity.

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1.2 Thesis Outline 7

The author dedicates chapter 4 to numerical optimization of magnetic actuatorsfor CubeSat applications. After outlining the state of the art of both actuatorsas well as their optimization, the formulation of a novel magnetic actuatoroptimization procedure is presented. Conclusively, results from the applicationof the proposed procedure are discussed.

Chapter 5 addresses the development of a novel CubeSat attitude controlactuator, the so-called picosatellite fluid-dynamic actuator (pFDA), by theauthor. Technological evolution of fluid actuators is revised first, followedby the derivation of design objectives for fluid actuator miniaturization. Toprepare the selection of the optimal fluid conduit geometry, angular momentumcalculation is derived for flat and three-dimensional conduits. Following asection dedicated to conduit design and manufacturing, miniaturization of thepump driver is addressed. Functional verification results in the estimation ofactuator dynamical properties. Based on the identified properties, a comparisonof reaction wheels and pFDAs is carried out in order to asses the agility ofthe two fundamentally different actuators. The final section is dedicated toredundancy concepts based on three-dimensional manifestations of conduits.

The work presented in chapters 3 to 5 is then used by the author in chapter 6 toresearch the implementation of a highly integrated, multi-functional solar panel.Starting out from the analysis and discussion of the current integration densityof CubeSat solar panels, a multi-functional solar panel design is addressed.Subsequently, assembly and test results are detailed before the chapter endswith the comparison between the realized solar panel and the criteria definedin section 2.3.

Finally, the author summarizes and concludes this thesis in chapter 7, andgives a recommendation for future work in the field of densely integratedCubeSat components and the application of pFDAs.

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2 High Performance Single Unit CubeSatDesign Approaches

Satellites in the CubeSat constellations addressed in section 1.1 rely oninexpensive commercial off-the-shelf (COTS) components, which becameavailable in recent years due to the technological advancement largely drivenby the consumer electronics market, as stated by Bandyopadhyay in [29].While Shimmin et al. in [30] state, that CubeSat subsystems built from COTScomponents have gained a high level of maturity, Selva and Krejci in [3] find,that scientific payloads are still limited by the well-known constraints for massand volume, as well as electric power and link budget. In comparison to tripleunit satellites, smaller CubeSats, especially below 2 U, were found to be rarelyused for "science missions with relatively high performance", as stated byPoghosyan and Golkar in [4].

2.1 CubeSat Market Observations

While conforming to the CubeSat design specification (CDS) [25], designand composition of present-day CubeSats greatly varies depending on thedeveloping/operating organization and their approach to satellite development.From the range of spacecraft addressed in section 1.1, three broader categoriesof CubeSats are derived:

Independent commercial CubeSats are completely designed and manufac-tured in-house by expert companies to meet the requirements of a singleconstellation (cf. section 2.1.1).

Commercially-procured CubeSats are used by organizations, that are moreinterested in flying their payload, than developing the spacecraft bus

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10 2 High Performance Single Unit CubeSat Design Approaches

(cf. section 2.1.2). Therefore, those organizations procure the completebus or the majority of the subsystems from CubeSat vendors.

Independent university CubeSats are developed by flagship or independentuniversities1, as defined by Swartwout and Jayne in [22] (cf. sec-tion 2.1.3). Those universities develop the complete satellite, and mayalso be responsible for the primary payload. This does not necessarilymean, that no parts of their CubeSats are procured, but they are ableto avoid parts that have negative influence to volumetric utilization orpower consumption on the spacecraft.

2.1.1 Independent Commercial Spacecraft

Probably the best example for independent commercial CubeSats are theDove satellites of Planet Labs’ Flock constellation (cf. page 2). They featurea telescope of about 2.5 U volume on a triple unit spacecraft. Due to thetelescope’s focal length, the image sensor is placed in a tuna can extension[25] at the rear end of the spacecraft. A narrow optical tube along the centeraxis of the satellite connects telescope and detector. Around this tube, thecomplete satellite bus was developed to be "a wrap-around design of totalvolume of about one-quarter of a Unit", as stated by Boshuizen et al. in [6].

This ultra-dense design was claimed by Boshuizen to be made possible usingtechnologies from the consumer electronics and automotive sector. Forattitude determination, the satellites feature a star camera, global positioningsystem (GPS), photo-diodes for coarse sun sensing, and a micro-electro-mechanical system (MEMS) inertial measurement unit (IMU). Attitude controlis enabled using a tetrahedral assembly of four reaction wheels and three air-core magnetorquers. Communication is based on a two-way UHF transceiver, S-band uplink receiver, and X-band downlink transmitter. A low-power processorequipped with solid state storage is running an unix-like operating system.Boshuizen et al. in [6] state, that none of these components "would have fitin a 0.5 U volume using currently available aerospace industry parts".

1For the sake of simplicity, the term independent is used here for both flagship andindependent universities to highlight the difference between a CubeSat bus developedcompletely by one university over a bus commercially-procured by an university.

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2.1 CubeSat Market Observations 11

2.1.2 Commercially-Procured Spacecraft

Unlike Planet Labs’s Dove satellites, Spire Global’s triple unit Lemur-2 Cube-Sats (cf. page 2) are no completely independent development of the company,but relying on the power system and solar cells commercially-procured fromClyde Space, as stated by Ostrove in [31].

Sky and Space launched the Three Diamonds demonstrator mission (cf. page 2,[8]). The satellites are based on the advanced triple unit CubeSat platform ofDanish supplier GomSpace [32]. The satellites of the Kepler Communicationsconstellation are based on the Clyde Space triple unit CubeSat bus [33].And Aerial & Maritime is procuring the first four satellites of its plannedconstellation from Danish CubeSat company GomSpace. The satellite bus isbased on the advanced GomSpace platform with heritage from the GOMX-3mission, which is presented by Gerhardt, Bisgaard, and Alminde in [34].

But not only NewSpace companies procure CubeSat buses on the market. Thespacecraft of NASA’s RAVAN mission as presented by Swartz et al. in [35]is based on Blue Canyon Technologies’s XB3 integrated triple unit CubeSatplatform [36]. RAVAN is a technology demonstration mission for radiometryinstruments used for analyzing the earth energy imbalance, which is a measurefor climate change (cf. Swartz et al. in [37]).

This enumeration of CubeSat missions based on commercially-procured space-craft buses does not claim to completeness. Most CubeSat suppliers likePumpkin, ISIS, SSTL, Clyde Space, or GomSpace are offering integrated plat-form solutions. To get a better understanding of those platforms’ capabilities,their characteristics have been collected in tables 2.1 to 2.3. The tables show,that the majority of integrated platforms aims at triple unit CubeSats, withonly three platforms aiming for single or double unit application. Besidesthe stated target CubeSat size, integrated platforms would be capable ofsupporting larger or smaller CubeSat form factors, with implications foremoston attitude control, due to the differing moment of inertia, and on payloadpower, due to solar cell area and therefore available power.

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2.1.2.1 Payload Volume

Single unit platforms allow for a maximum payload volume ratio of 25 %. Incomparison to the larger platforms, which are in a range between 50–66.7 %of payload volume, this appears to be very small. The observed behavior isexplained by the fact, that single and triple unit platforms of the same vendorrely on a similar set of CubeSat subsystems. For compatibility reasons withother suppliers, those subsystems are implemented on individual printed circuitboards (PCBs) and conform to the PC/104 form factor, which is the de-factostandard used for board-to-board connectors in the CubeSat industry [38].

2.1.2.2 Payload Mass

While the CubeSat design specification [25] intends for a total of 1.33 kgmass per unit, spacecraft like Planet Labs’s Dove [6] or NASA’s RAVAN [46]weigh more than the 4 kg allowed for a triple unit CubeSat. Other vendors ofCubeSat deployers, however, have establisehd a 2 kg per unit maximum masslimitation. The same limit has been adopted by integrated platform vendors (cf.table 2.1). Hence, overall satellite volume translates to a total maximum mass

Table 2.1: Integrated CubeSat platforms: mechanical and electrical characteristics

Brand Name Satellite Payload SourceVolume Volume Mass Power

U U kg WClyde Space 1 0.2 2 [33]GomSpace Basic 1 0.25 0.17 1.3 [32, 39]Space Inventor 2 1.25 0.3 20 [40, 41]BCT XB3 3 2 4.5 60 [42]Clyde Space 3 1.6 12 [33]GomSpace Advanced 3 2 3.5 [32]ISIS Basic 3 2 4 2 [43]ISIS Advanced 3 1.5 3 3.5 [43]SSTL Cube-X 3 2 2 6 [44]Tyvak Endeavour 3 2 65 [45]

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of 2 kg, 4 kg and 6 kg for single, double, and triple unit CubeSats, respectively.Based on available payload masses and the above maximum masses, singleand double unit CubeSats achieve a rather homogeneous payload mass ratioof about 8 %. The very small payload mass ratio of the two-unit platform,in contrast to its 62.5 % payload volume ratio, is explained by the aluminumhousings used for all subsystem to increase radiation tolerance [40], makingthem significantly heavier than the subsystems of most other companies.

Triple unit payload mass ratios show a wide spread, spanning from 33.3 % forthe Cube-X [44] to 75 % for the XB3 platform. Due to poor documentationin the original sources, no clear trend for payload mass ratio is distinguishable.However, available payload mass of 3 kg for a triple unit platform is realistic,as three out of four entries show 3 kg and above available payload mass.

2.1.2.3 Payload Power

In their 2010 survey [47], Bouwmeester and Guo state, that for CubeSats withbody-mounted solar cells the available specific power increases for smallersatellite masses. They explain this with the fact, that satellite mass is relatedto the outer satellite dimensions with the third power, while available solarcell area is related only with the second power to the outer dimensions. Forsatellites with a mass between 1–2 kg, Bouwmeester and Guo document amaximum specific power of 2 W/kg. It drops to about 1 W/kg for satellitesin the range of 4–6 kg.

Table 2.1 lists a maximum payload power of 2 W/U for the single unit plat-forms, equal to a specific payload power in the range of 1–2 W/kg, whichagrees well with the specific powers presented by Bouwmeester and Guo in[47]. The situation for double and triple unit spacecraft appears to be veryinhomogeneous, with available payload powers between less than 1 W/U and21.7 W/U. Expressed as specific power, this would be in a range of up toabout 16 W/kg.

Observed available payload power of single unit spacecraft is well explained bythe fact, that the small satellites usually do not feature deployable solar panels.Values of 4 W/U and above seen for the double and triple unit variants are

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also well explained with the application of deployable solar panels. Smallervalues of less than 4 W/U were observed for several platforms. A closer lookat the original data sources shows, that except for the ISIS basic bus, thegiven average payload powers are the absolute minimum values, and mayeasily surpass 24 W, as stated e.g. in [43].

2.1.2.4 Attitude Determination

According to the survey of Bouwmeester and Guo [47], sun sensors andmagnetometers are the most commonly used sensor type within all pico- andnanosatellites between 1957 and 2009. Nearly 30 % of all missions in theirdatabase were equipped with either one or both of these sensor types. Funkeet al. in their 2016 analysis on the characteristics and development of smallsatellites [48] state, that coarse sun sensors, magnetic field sensors, and angularrate gyroscopes have become the state of the art for small satellite attitudedetermination. To achieve precision attitude knowledge, however, additionalstar cameras are necessary.

The column for attitude knowledge data gathered in table 2.2 shows thebiggest gaps. Only three triple unit platform manufacturers explicitly statethe attitude knowledge. Tyvak’s is ranked top among the others with fewarcsec attitude knowledge, while Clyde Space’s and GomSpace’s platformssupport attitude knowledge in the range of few arcmin.

2.1.2.5 Attitude Control

In table 2.2, a clear distinction can be seen between low-precision systemswith 5° and above pointing accuracy, and high-precision systems that featurea pointing accuracy of better than 1°. A similar distinction is made for agility,represented by the slew rate of maneuvers. While GomSpace’s single unitCubeSat platform has only a slew rate of 0.17 °/s, all triple unit platformsfeature 3 °/s and above. The larger slew rates of the triple unit platforms areachieved using sets of at least three reaction wheels.

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With a pointing accuracy of 7.2 arcsec and a slew rate of 10 °/s, Blue CanyonTechnologies’s integrated triple unit CubeSat platform stands out from allother platforms. This performance is achieved using reaction wheels and asensor bank that comprises, among other, two star cameras.

2.1.2.6 Communications

According to Funke et al. in [48], the most commonly used band for com-munication on CubeSats is UHF. All basic configurations in table 2.3 hostUHF communications equipment, and achieve up to 19.2 kbps data rate. Themajority of advanced configurations feature at least S-band communication,which starts at a data rate of 3.40 Mbps. The high-end variants offer additionalX-band communication, with data rates up to 800 Mbps in the case of theClyde Space platform.

2.1.2.7 Harness and Connectors

Bouwmeester, Langer, and Gill in [38] state, that the PC/104 connector"has become the de-facto standard wiring harness in CubeSats, as mostcommercial developers provide their subsystems with this interface". Theirsurvey showed, that a growing number of CubeSats to be launched willimplement the PC/104 connector, which they justify with the increasingavailability of commercial subsystems that feature the 104-pin stackableconnector. They further conclude, that a small majority of participants intheir survey states that the PC/104 connector is too big, while otherwise nomajor problems are seen with the standard.

Bouwmeester, Langer, and Gill in [38] recommend that a future standardinterface connector for CubeSat subsystems should be smaller than the PC/104connector. They further conclude that this standard interface should have afixed pin allocation to achieve general compatibility between subsystems.

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Table 2.2: Integrated CubeSat platforms: attitude determination and control systemcharacteristics

Brand Satellite ADCS SourceVolume Knowledge Accuracy Slew

U arcmin ° °/sClyde Space 1 5 [33]GomSpace 1 0.17 [32, 39]Space Inventor 2BCT 3 0.002 10 [42]Clyde Space 3 3.5 5 [33]GomSpace 3 6 0.1 [32]ISIS 3 10 4 [43]ISIS 3 1 3 [43]SSTL 3Tyvak 3 0.043 0.057 3 [45]

Table 2.3: Integrated CubeSat platforms: communication system characteristics

Brand Satellite Communication SourceVolume Band Rate

U kbpsClyde Space 1 VHF, UHF 9.6 [33]GomSpace 1 UHF [32, 39]Space Inventor 2 UHF 19.2 [40]BCT 3 UHF, S, X 15 000 [42]Clyde Space 3 VHF, UHF, S, X 800 000 [33]GomSpace 3 UHF, S 7 500 [32]ISIS 3 VHF, UHF 9.6 [43]ISIS 3 VHF, UHF, S 3 400 [43]SSTL 3Tyvak 3

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2.1.3 Independent University Spacecraft

Flagship and prolific independent universities are the groups that dominate theeducational CubeSat launches, as stated by Swartwout and Jayne in [22]. Thefirst flagship university listed in their list of spacefaring universities happens tobe TU Berlin. Following the 2016 launch of BEESAT-4 [49], the 2017 launchof TechnoSat [50], and the 2018 launch of the S-Net formation [23], TUBerlin has launched a total of 16 small satellites over the last 27 years. Withthe upcoming launches of the BEESAT-5 to BEESAT-8 [26] formation andTUBIN [51] over the next two years, the number of small satellites launchedby TU Berlin will increase to 21.

2.1.3.1 TU Berlin Experimental and Educational Satellites

The most recent CubeSat mission of TU Berlin is the single unit spacecraftBEESAT-4. Its design is based on the BEESAT and BEESAT-2 missions andthe satellite was deployed from the larger, German Aerospace Center (DLR)-operated, BIROS spacecraft after four months in orbit. Primary missionobjective is the in-orbit demonstration of the Phoenix nanosatellite GPSreceiver [52]. Most of the information on the BEESAT-4 satellite in thissection is taken from the 2017 publication of Weiß and Kapitola [53].

Overview of BEESAT-4

BEESAT-4 features precision sun sensors and MEMS magnetic field andangular rate sensors for attitude determination. Three RW 1 reaction wheelsare used for three-axis attitude stabilization. Six magnetic coils support wheelangular momentum desaturation and spin dampening using a �̇� control law.

The only subsystem not developed at or in cooperation with TU Berlin is theBK77 UHF transceiver used for communication. On-board computer (OBC),power control and distribution unit (PCDU), and payload data handling (PDH)boards were all developed by TU Berlin. The objective during satellite busdevelopment was to achieve a complete as possible single-failure tolerance ofevery subsystem and therefore the complete satellite [54].

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Electrical connection between the subsystems is not based on the PC/104standard, a custom COTS connector solution is used. Communication betweensubsystems is happening on two independent controller area network (CAN)buses. The solar panels, which feature magnetorquers and TU Berlin’s precisionsun sensors, are connected to the power control unit (PCU) board withinsulation-displacement contact (IDC) connectors and flat ribbon cables [54].

All subsystems except for the PDH board, have flight heritage from BEESATand BEESAT-2. The interior of the satellite is divided in two by the aluminumbattery compartment. OBC, PCDU, and PDH boards occupy one half of thesatellite. In the other half, the reaction wheels, wheel drive electronic (WDE)board, and two BK77 UHF transceivers are located. The two deployableantennas are also attached on this side of the spacecraft. Six solar panelscomplete the satellite, each hosting multiple solar cells, a sun sensor, and amagnetorquer. The panel close to the PDH hosts the GPS antenna. This oneand the panel on the side of the deployer’s access panel use 20 × 20 mm cells,all other panels are identical and feature two 80 × 40 mm cells each.

The primary payload of the satellite is a GPS receiver, complemented by acamera as secondary payload. Both payloads are mounted to the PDH board.From BEESAT-2 to BEESAT-4, the PDH board needed to be redesigned [54],as the GPS receiver requires a 5 V power supply, which was not available since.Payload volume is approximated to be 0.125 U, and the payload mass ratioof the 1 kg satellite is 7.18 %. Maximum power consumption of the PDHis 1.4 W [55]. However, not all payloads are operated simultaneously, andnominal average payload power for typical operational scenarios might be less.

BEESAT-4 in Comparison

BEESAT-4 overall payload volume ratio of 12.5 % appears to be very small incomparison to other single unit platforms (cf. table 2.1), which allow for aratio of 20–25 %. Owing to its heritage from BEESAT and BEESAT-2, thesmall payload volume ratio takes no wonder. For BEESAT, where the payloadconsisted of the reaction wheel assembly (RWA), WDE, and the PDH boardincluding the camera, the estimated payload volume ratio is about 35–40 %.

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Comparing BEESAT-4’s mass of about 1 kg to the de-facto upper limit of 2 kg(cf. section 2.1.2.2), the satellite bus would already allow for a payload massratio of about 53.6 %. Given a CubeSat with homogeneous mass distribution,moment of inertia 𝐼 about one primary axis is given as

𝐼 = 𝑀

6 𝑙2, (2.1)

where 𝑀 is satellite mass and 𝑙 is the edge length of the satellite. Doublingthe mass from 1 kg to 2 kg would therefore lead to double the inertia, cuttingthe maximum achievable angular rate in half. On-orbit results for attitudecontrol data presented by Weiß and Kapitola in [53] document, that thespacecraft performs slew maneuvers at an angular rate of 5 °/s. During thismaneuver, the reaction wheels reach a speed of about 2 500 rpm, which is onlyabout 31.3 % of the nominal, and 15.6 % of the maximum wheel speed. Thisunderlines, that the existing BEESAT satellite bus would be able to host heavierpayloads. Weiß and Kapitola in [53] state, that the active attitude controloperational time is currently limited by the power budget to a maximum of60 min. Therefore the power budget would be the limiting factor to increasingthe overall mass of the existing satellite.

Comparison of available payload power is more difficult to conduct. Maximumpayload power consumption stated on page 18 is not representative for averagepower consumption at nominal satellite operation with GPS and inter-satellitelink (ISL) experiments (cf. [53]).

With an attitude error of about 5° based on sun sensor and magnetic fieldmeasurements, BEESAT-4 is comparable to basic CubeSat platforms (cf.table 2.2). Slew rates of 5 °/s are positioning BEESAT-4’s agility among thehigher performing triple unit platforms.

Nominal data rate of BEESAT-4 is 4.8 kbps, with the option to use 9.60 kbps,according to [54]. This places the BEESAT-4 communication system atthe lower end of the integrated platforms (cf. table 2.3). Therefore, thecommunication system is a bottleneck for possible high performance payloadson the existing BEESAT bus.

The BEESAT satellites do not rely on subsystems that conform to the PC/104standard. This offers flexibility in terms of harness and connectors, and allows

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to put the reaction wheels on a common baseplate in one sixth of the volumeof the satellite, while the WDE and two UHF transceivers are occupyingthe adjacent sixth of the internal volume. An assembly using subsystemsconforming to PC/104, like e.g. the attitude determination and controlsystem (ADCS) of GOMX-3 which uses a tetrahedron RWA based on fourRW 1 reaction wheels and the associated WDE board mounted on a largermotherboard [34], utilizes more satellite volume.

2.1.3.2 Universität Würzburg Experimental Satellites

A second alternative to using PC/104 compliant subsystems is pursued by theUniversität Würzburg Experimentalsatellit 3 (UWE-3). Busch in his disserta-tion on robust, flexible, and efficient design for miniature satellite systems [56]states, that "experiences from the previous UWE missions indicated that there-utilization of a system, which is not inherently designed to be extended, canhardly be upgraded without producing a significant increase in total systemcomplexity". For the design philosophy of UWE-3 and his successor missions,this lead to a preservation of the authority over the subsystem interface, thusrelying on third party products only on the component level. The followingdiscussion of the UWE-3 bus is mainly based on the dissertation of Busch [56]and his related publications.

Using a backplane, the UWE-3 bus design avoids the need for cables. Allsubsystem boards are directly connected to the backplane using standardizedconnectors. This connector implements the subsystem interface, which com-bines power lines for different voltages, GPIO, and digital interfaces like UARTor I2C. In addition, dedicated lines for global reset and debug support arelocated on the electrical bus. Malfunctioning subsystems may be completelypowered down to protect the digital buses from undesired bus allocation.Umbilical connections and the remove before flight (RBF) pin are realizedusing the first board connected to the backplane, the so-called front accessboard (FAB), which holds the required external connectors. All solar panelsare attached directly to either the backplane or the FAB [57].

While the backplane and the FAB have advantages in terms of defined interfacesas well as easy access and testability, they are very sparsely populated with

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electric components and thus occupy volume, that could be used by othersubsystem boards or payloads. Also the support of multiple digital interfaceslike universal asynchronous receiver transmitter (UART) or inter-integratedcircuit (I2C), of which some were not designed to be used in harsh environmentsand therefore drive the need for additional components for selective busisolation, is criticized in comparison to the use of a single, robust interfacethat was designed for harsh environments, like e.g. CAN bus. The use of abackplane has the further disadvantage that late changes to the arrangementof subsystems are rendered impossible owing to the need of having a newbackplane PCB manufactured and equipped with connectors.

UWE-3 uses spacers in the board stack together with four identical rails asbase for the mechanical structure. The solar panels, consisting of a two-layerPCB with an aluminum core, complete the satellite structure. Busch claims,that the panels provide "functionalities of mechanical stability, thermal balance,radiation protection, antenna ground plane, and power or sensor electronics inan integrated compact and lightweight design" [56].

Due to the two-layer stack-up and the aluminum core, it is difficult to increasethe component count on the UWE-3 solar panels. In particular as the air-coredcoils further limit the available surface area on the panels.

The UWE-3 bus, in conclusion, offers major advantages over existing CubeSatbuses in terms of ease of assembly, subsystem interfacing, and robustness. Yet,there is seen potential for further advancing the bus concept and miniaturizingthe volume required for the backplane and FAB.

2.1.3.3 Miniaturized Components for Distributed Pico- andNanosatellite Systems

In the scope of the PiNaSys and PiNaSys II projects at TU Berlin, researchersaim at further miniaturizing CubeSat components to enable distributed satellitesystems. Baumann et al. are developing a swarm consisting of four quarter-unit CubeSats, named BEESAT-5 to BEESAT-8, that will be deployed togetherfrom a single unit CubeSat deployment container in 2018 [28].

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To miniaturize the satellite to one quarter of an unit, all satellite subsystemsare integrated on one PCB, except for the batteries, which are mounted insidethe metallic structure of the satellites close to the center of mass. Almostcomplete single point of failure tolerance is achieved by using the same boardtwice, and connecting them via a redundant CAN bus. The primary missionconcept is to demonstrate the miniaturized bus, which is equipped with anumber of novel, highly miniaturized components for communication andattitude determination, as stated by Baumann et al. in [26].

The PCBs of BEESAT-5 to BEESAT-8 feature a very high level of integration,as PCDU, UHF communication equipment, OBC, ADCS, the optical payload,and the experimental X-band transmitter are all integrated on on side of thepanel. Internal communication between those subsystems is organized via aredundant CAN bus that is managed by the PCU [26]. The other side of thePCBs hosts a 80 × 80 mm solar cell and all antennas.

Being close to the total-integration and ultra-high density approach demon-strated by Planet Labs’ Dove satellites, the chosen architecture increasescomplexity during development, as multiple engineers need to work on thesame PCB design to integrate their subsystems. Development cost for suc-cessor missions with different mission objectives is increased due to the largeamount of work for PCB design changes.

As the complete satellite is an in-orbit demonstration (IOD) mission, it isdifficult to define a ratio between payload and bus volume and comparethis to commercially procured CubeSats based on integrated platforms (cf.section 2.1.2). From the subsystem point of view, the BEESAT-5 to BEESAT-8satellites will demonstrate high performance attitude and orbit determinationcapabilities using the experimental star cameras and GPS receivers. Asmagnetorquers are the only means for attitude control, precise and agilepointing is out of range. The X-band transmitter provides payload datadownlink with a rate of up to 1 Mbps using a transmitting power of 400 mW.Compared to X-band downlink data rates of up to 800 Mbps (cf. table 2.3),this is very small and better comparable to communication in the S-band.However, other communication systems require much higher amounts ofelectric energy to achieve the high data rates.

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2.2 High Performance Single Unit CubeSat Design Criteria 23

2.2 High Performance Single Unit CubeSat DesignCriteria

From the discussion of the characteristics of commercially-procured andindependent university spacecraft in sections 2.1.2 and 2.1.3, respectively, acertain set of design criteria for a high performance single unit CubeSat arederived. Compatibility with existing solutions offered on the CubeSat marketis not a requirement.

2.2.1 Provided Payload Resources

The ratio between available payload volume and overall CubeSat volume doesnot scale for single unit CubeSats in comparison to the larger CubeSats: Whiletriple unit spacecraft have at least 50 % of volume available for payloads, singleunit CubeSats allow only for up to 25 % (cf. section 2.1.2.1). Main reasonfor this is the utilization of CubeSat subsystems that obey to the PC/104standard (cf. [38]). Avoiding PC/104-based subsystems and exploiting thelevel of miniaturization seen in the scope of the PiNaSys and PiNaSys IIprojects, a better volumetric utilization of about 50 % with available payloadmasses as high as 1 kg should be achievable.

Section 2.1.2.3 documents 2 W average payload power available on single unitplatforms. A general increase in bus performance leads to an increased buspower demand. Keeping 2 W power available for the payload therefore mightrequire to use a set of deployable solar panels on a single unit CubeSat.

2.2.2 Attitude Determination and Control Capabilities

Simple platforms achieve attitude knowledge in the range of few arcmin,while advanced platforms employ star cameras to reach few arcsec. Forpointing accuracy, a distinction is made between low-precision platformswith a maximum accuracy of 5° and above on the one hand, and high-precision platforms with a maximum of 1° and below on the other hand.High-precision platforms are possible utilizing sun sensors in combination with

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MEMS magnetic field and angular rate sensors for attitude determination andreaction wheels with magnetorquers for attitude control. Reaction wheelsallow for larger slew angles and the availability of control torque about allthree axis at all times, in comparison to magnetic-only attitude control.

2.2.3 Downlink Capabilities

Analysis of downlink data rates of single and double unit CubeSat platformsreveals, that they currently rely on UHF communication with data rates up to19.2 kbps (cf. table 2.3). The majority of triple unit spacecraft feature S-bandwith data rates starting from 3 400 kbps. However, no data on power demandof the featured communication hardware is provided. For a power-limitedsingle unit CubeSat, the use of a receiver like HISPICO [58, 59] is suggested,as it provides up to 1 Mbps of downlink data rate and is specifically designedfor CubeSat applications. Utilization of a S-band transmitter drives the needfor a patch antenna on the satellite.

2.2.4 Data and Power Bus

Following the recommendations of Bouwmeester, Langer, and Gill (cf. sec-tion 2.1.2.7), data and power bus should not be implemented based on thePC/104 standard. A backplane featuring identical interface connectors forevery subsystem, as used on UWE-3 and its successor missions (cf. sec-tion 2.1.3.2), is already an improvement in comparison to PC/104. How-ever, backplanes and large subsystem connectors are negatively affecting thepayload-to-bus ratios. While using a backplane, the modular TUBiX20 plat-form developed at TU Berlin offers a common data and power bus whereall inter-subsystem communication is carried out on a redundant CAN businterface (cf. [60]). This approach to modular satellite architecture allowsto adapt the platform to new missions with a minimum effort in terms ofredesign of existing hardware, while it allows for full flexibility in adding orremoving functionality.

The CubeSats of TU Berlin, like BEESAT-4 and the upcoming BEESAT-5 toBEESAT-8 satellites, already feature several subsystems, that are connected

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2.3 Highly Integrated Solar Panel Design Criteria 25

to a common, redundant CAN bus. For BEESAT-4 and his predecessors,modularity was no key design criteria, which lead to a number of redesigns ofthe PDH board (cf. [54], page 18). The BEESAT-5 to BEESAT-8 satellites,which implement the complete satellite electronics on two identical PCBs, arealso no modular designs and extension for future missions will require a lot ofwork for redesign.

To overcome the aforementioned limitations, a common data and power buswith a standardized connector interface, that does not require a backplane isrequired. One possible implementation could be based on IDC connectors andflat ribbon cable.

2.3 Highly Integrated Solar Panel Design Criteria

To support the criteria for high performance single unit CubeSat design derivedin sections 2.2.1 to 2.2.4, subsystem components can be relocated to thesolar panel PCBs, which are currently not very densely populated with electriccomponents in comparison to the subsystem boards.

Moving components from the subsystem boards in the main PCB stack andintegrating multiple subsystems on single boards in the center of the satellitefrees up payload volume. Moreover, the mass of the reduced PCB substratebecomes available for the payload. If the solar panels integrate maximumpower point tracker (MPPT) capabilities directly on the panel, additionalpayload power is freed up. The mentioned measures all support the criteriafor providing increased payload resources laid out in section 2.2.1.

Modern CubeSat solar panels (cf. section 6.1) already support a certainamount of the attitude determination and control capabilities addressed insection 2.2.2. Consequently, sensors and actuators for attitude determinationand control should all be moved to the solar panels, to support the criteria forincreased payload resources stated in section 2.2.1. Implementation of attitudesensors, magnetorquers, and angular momentum exchange-based actuators onthe same solar panel together with a microcontroller allows to locally filterattitude determination data, close the control loop, and provide reaction wheelangular momentum desaturation.

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While star trackers have undergone a significant miniaturization over the lastdecade and are now available for triple unit CubeSat platforms, they are notforseen for integration on a multi-functional solar panel in this work. Thisis mainly due to the geometry and volume occupied by even the tiniest startrackers: Due to the shape of the baffle, star trackers would stick too far intothe satellite bus to design a highly-integrated solar panel that supports a large,compact payload volume on the satellite.

Addressing the criteria for payload data downlink capabilities derived in sec-tion 2.2.3, solar antennas offer the advantages of having a S-band patchantenna available while counteracting the negative effect of reduced availablesolar cell area on the surface on the satellite (cf. chapter 3). While a relocationof the S-band transmitter components would positively influence the criteriadefined in section 2.2.1, data transfer from the PDH board to the solar panelis limited to a maximum of 1 Mbps on the CAN bus. A direct wire connectionbetween the solar antenna and the transmitter in combination with a highthroughput interface between transmitter and PDH is required to obtain highdownlink rates beyond 1 Mbps from orbit.

The components integrated on the solar panel, which traditionally belongto multiple subsystems, connect to the same standardized data and powerbus as all the other subsystems (cf. section 2.2.4). If no voltage regulatorsare located on the solar panel to locally provide power to the subsystemcomponents, dedicated power supply lines that feed the panel need to bepresent on the bus. Dedicated unregulated power lines are required to feedthe electric power generated by the solar cells to the batteries that are locatedin the central part of the CubeSat bus.

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3 Picosatellite Solar Antenna

One bottleneck for CubeSat missions is limited downlink data rate due tocommunication in the UHF/VHF bands. Those bands are commonly usedon CubeSats as they do not require precise pointing control, and deploymentmechanisms for monopole antennas are simple and therefore robust andreliable, as stated by Funke et al. in [48]. Bouwmeester and Guo in [47]state, that downlink data rates in the UHF band are commonly in the rangeof 1.2–9.6 kbps.

Increasing downlink data rates requires the use of higher frequency bands.S-band is a viable solution that enables several Mbps of bandwidth (cf. [61],table 2.3). In order to realize high gains, S-band antennas require very narrowbeam widths and may be realized as circular polarized patch antennas, asstated by Kakoyiannis and Constantinou in [62]. Apart from increased pointingdemand, application of patch antennas reduces the available solar cell area onCubeSats.

So-called solar antennas, which integrate antennas and solar cells, are oneway to mitigate this loss in solar cell area. As the electromagnetic propertiesof the solar cell might interfere with the radiation pattern of the antenna,this influence needs to be investigated. The development of a S-band solarpatch antenna for application on a multi-functional CubeSat solar panel wascarried out in the scope of the MKI project by researchers at Fraunhofer IZMin cooperation with the author. This chapter discusses this research work,published in parts in [63, 64, 65] under the direction of the author.

3.1 Solar Antenna State of the Art

Extensive discussion is available on all possible antenna types used for smallsatellite missions in [62]. In the scope of the present work, only planar

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28 3 Picosatellite Solar Antenna

solar antennas are considered due to size constraints and the want to avoidadditional mechanisms to release antenna structures.

Tanaka et al. in [66] were the first to mention solar antennas for small satelliteapplications. The concept presented there places the solar cells on top of apatch antenna. A patch antenna, being a lossy cavity, radiates via the fringingfields along the open edges of the patch. As long as the solar cell is placedin such a manner that the radiation pattern of the underlying patch is notdisturbed and the cell is able to generate power, this was found to be a validcombination to form a solar antenna.

The replacement of the radiating patch element in a planar antenna by a solarcell is presented by Henze et al. in [67, 68, 69]. In order to avoid harmfulinterference between antenna and solar cell, additional electric componentsdecoupling the radio frequency (RF) part from the direct current (DC) partare required.

Based on amorphous silicon solar cells, Vaccaro et al. in [70] describe thedevelopment of circularly polarized slot antennas embedded within solar cells.The thin and flexible cells feature a crossed-slot antenna for circular polarization.A first in-orbit verification mission using advanced solar antennas was launchedin the year 2005, as stated by Markgraf et al. in [71]. One integrated antennawas designed for GPS signal reception and the other one for S-band downlink.The experiment worked for over two years on orbit, and the promising flightresults were presented by Vaccaro et al. in [72] in 2009. However, instead ofusing amorphous solar cells, high-efficiency GaAs cells were used.

Printing meshed patch antennas allows to reverse the order in the stack-upof antenna and solar cell, as meshed patches are transparent for the incidentlight, but not for the RF waves. Related work is presented by Turpin andBaktur in [73], as well as by Yasin and Baktur in [74]. They found, that forthis type of antenna the interaction between solar cell substrate and antennaperformance is very pronounced.

An in-depth discussion on planar solar antenna state of the art was alreadypresented in [63]. A lot of ongoing research on solar antennas was found inthe literature. However, only the work of Vaccaro and Markgraf lead to actualon-orbit verification (OOV) of solar antennas [71, 72]. The two antennas

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3.2 Patch Antenna Development 29

presented there demonstrated an on-orbit lifetime of over two years. To theknowledge of the author, however, no solar antenna so far has found its wayinto application on small CubeSat missions.

3.2 Patch Antenna Development

Development of a solar antenna requires development of a S-band patchantenna first. The radiation mechanism of a solar patch antenna is shown infigure 3.1. Patch and ground plane form a lossy cavity that acts as a resonatorfor alternating electromagnetic fields. Energy is released from the resonatorin form of fringing fields at the edges of the antenna patch. The emittedelectromagnetic field is not passing through the patch. If the solar cell edgesare recessed enough from the active patch edges, electromagnetic interferencebetween solar cell and antenna is reduced. The setup shown in figure 3.1 isthat of a linearly polarized antenna. Circular polarization is achieved by tuningthe patch shape away from a perfect square and putting the feed probe atthe optimal location with respect to the patch.

The patch antenna is designed against the requirements of the S-band trans-mitter HISPICO, which was developed in a joint project under the lead of TUi

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ground plane

patch

solar cell

fringingfield

fringingfield

voltage

current

Figure 3.1: Radiation mechanism of a patch antenna and integration of a solar cell(adapted from [63])

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Berlin (cf. Alavi and Rießelmann, [58]). Its small size and power demand allowthe application on single unit CubeSat missions. The transmitter operatesin the range of 2.2–2.3 GHz, the channel is adjustable in steps of 100 kHz,and it allows for up to 1.06 Mbps user data rate. It requires an antennawith a gain of approximately 6 dB and a 3 dB RF bandwidth of 50 MHz [59].Those parameters define the geometric dimensions of the antenna patch, andtherefore its mass and the maximum size of the solar cell on top of it.

Simulation results of six different patch shapes and three different substratethicknesses were presented by Grau et al. in [63]. To find the thickness bestmatching the required impedance bandwidth of the antenna, return loss of aquasi-square patch was simulated for all three thicknesses (cf. figure 3.2). Onlythe patch antenna with maximum thickness fulfills the impedance bandwidthrequirement over the full transmitter frequency range. Thinner substrateshave impedance bandwidths of less than 50 MHz. Therefore, they are notapplicable over the full frequency range. Using thinner substrates would requireto adapt the antenna patch to a predefined transmission channel. If frequencycoordination happens in a late project phase, using thinner substrates couldlead to problems with respect to mission time and cost budgets. Thickersubstrates require more care in mounting the antenna to the satellite sidepanel. Due to their increased substrate mass, mechanical loads during rocketlaunch introduce more stress to the locking mechanism that keeps the antenna

Figure 3.2: Return loss and axial ratio of a quasi-square patch antenna for differentthicknesses of the dielectric medium [63]

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3.2 Patch Antenna Development 31

in place. As it was decided to use electrically conducting adhesive to connectthe solar antenna to the satellite side panel, the area on which the adhesiveis applied needs to be large enough to withstand the vibrational and shockloads during rocket launch.

After finding the required thickness of the quasi-square patch, other patchshapes were simulated and compared to each other for the maximum thickness(cf. [63]). Results show, that there is no significant difference in return lossand axial ratio for the different shapes (cf. figure 3.3). The solid line thatrepresents the quasi-square patch, however, shows the best performance ofall patch shapes and this patch shape was therefore selected to form thefoundation of the solar antenna.

Figure 3.3: Return loss and axial ratio of all researched patch antennas [63]

Simulation results for all six patch shapes show that right hand circularpolarization (RHCP) gain is about 6.50 dB, and less than −20 dB for lefthand circular polarization (LHCP) in the main beam direction (cf. figure 3.4).For RHCP, the differences between all beam shapes in forward direction arenegligible. The backward radiation differs more but stays well below −10 dBin opposite beam direction. Half power beamwidth (HPBW) in this case isabout 85°.

Similar behavior is observed for RHCP and LHCP gains in the yz plane (cf.figure 3.5). Differences between the patch shapes are even less articulated,

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Figure 3.4: Right hand and left hand circular polarization antenna gains for allresearched patch geometries in the xz plane in dB [63]

Figure 3.5: Right hand and left hand circular polarization antenna gains for allresearched patch geometries in the yz plane in dB [63]

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3.3 Solar Cell Integration 33

and gains are all in the same ranges as for the xz plane. Here as well, HPBWis about 85°.

Simulations have shown, that an antenna using a quasi-square patch is a validoption to realize a solar antenna respecting the requirements stated at thebeginning of this section. Thinner substrates could only be used for antennas,that would not fulfill the requirements over the complete transmitter frequencyrange. The large mass required due to the therefore required thicker substrateneeds to be compensated by a large contact area between antenna and solarpanel, if some sort of adhesive is used for mounting the antenna.

3.3 Solar Cell Integration

The in-plane substrate dimensions used during simulations are too large tofit the solar antenna onto the CubeSat panel, so the substrate was resizedto 42 × 40 × 5 mm. Simulations showed that the increase in thickness isnecessary in order to achieve the required RF characteristics. Additionalcopper layers were added above the antenna patch for the same reason. Theantenna mounted on the multi-functional solar panel (1) is shown in figure 3.6.Antenna substrate (3) features a copper ground plane (2) on the bottom andembeds the patch (4) and beam forming wedges (7). A single probe (9) isused to feed the RF signal to the antenna patch through the miniature SMAi

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1

2

3

4

56

7

8

9

10

Figure 3.6: Cross section schematic of solar antenna (adapted from [65])

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connector (10). On top of the substrate, additional copper areas and traces(5) are used to electrically interconnect the solar cell (6). The fringing fields(8) are forming at the patch edges.

The fully assembled solar antenna is shown in figure 3.7. The 20 × 20 mmsolar cell (6) is mounted on top of the dielectric (3) using conductive adhesive.The border around the cell is solder resist, used to define the area in which toapply the adhesive. From the copper patch beneath, one track runs directlyto the edge of the antenna, where a cut-through via (11) is used to routeelectric power from the top to the bottom of the antenna. Bond wires areused to connect the top connector (12) of the solar cell to the copper track(13) that routes the electric power from the cell to the bottom of the antenna.The cut-through holes end in half-circular pads on the bottom of the solarantenna. They are used to solder the antenna to the solar panel in order tofeed the electric power to the MPPT circuitry located on the solar panel (cf.section 6.2). Pads and antenna ground plane are separated, so that no directelectrical connection is made between the two. Beam forming wedges (7) arevisible through the substrate top layer (3).

3

76

11

1213

Figure 3.7: Image of the fully integrated solar patch antenna [65]

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3.4 Solar Antenna Functional Verification 35

With the solar cell and the connector integrated as shown in figure 3.7, solarantenna total mass is less than 18 g. Due to the bond wires, additional careneeds to be taken while handling the solar antenna during assembly, integration,and test. Wire-bonding also increases the solar antenna manufacturing costs.This does not only include bonding itself, but also the plating on the PCB.

3.4 Solar Antenna Functional Verification

During solar antenna development, functional verification was demonstrated.Environmental tests are scheduled on system level, in this case for the fullyintegrated solar panel (cf. section 6.3.3). Verification campaign was conductedat Fraunhofer IZM and results first published in [65] under the direction of theauthor. To verify the solar antenna’s RF characteristics, a load resistor wassoldered in between the vias of the antenna. Return loss and the antenna gainwere measured while a light source was applied to the solar cell. Measuredreturn loss is shown in figure 3.8.

In Comparison to the simulation results (cf. figure 3.3), measurements show areduction in impedance bandwidth of 10 MHz. Minimum return loss frequencyshifts towards the lower limit of the targeted frequency band, and the −10 dBreturn loss criterion is violated in the range of 2.24–2.27 GHz. It is unknown,whether these changes are induced by solar cell integration, or by reducedsubstrate size, as no separate measurements without solar cells were performed.

Figure 3.8: Solar antenna return loss measurement [65]

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During simulations, substrate area was assumed to be larger, and the antennapatch was aligned under 45° with respect to the substrate edges.

Comparison of solar antenna RHCP gain measurements (cf. figure 3.9) tosimulation results (cf. figures 3.4 and 3.5) shows that the angle of the halfpower beamwidth is reduced to about 75°. Gain shows a lobular structure,and a pronounced asymmetry between the left and the right half of the plot.Asymmetry is introduced by the copper track used to electrically contactthe solar cell to the underlying printed circuit board of the solar panel (cf.figure 3.7). Antenna gain in main beam direction is still above 5 dB.

This change in antenna radiation pattern increases requirements for the ADCSof a CubeSat which is equipped with the developed solar antenna. To realizea high downlink data rate, such a spacecraft would need to point more precisetowards the receiving ground station antenna. This, in fact, increases therequired attitude knowledge and pointing accuracy capabilities of the satellitebus (cf. sections 2.1.2.4 and 2.1.2.5).

Figure 3.9: Solar antenna gain measurement [65]

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4 Magnetic Actuator Optimization

Magnetic actuators provide the most basic means for attitude control. Theyare often used in pico satellites, because they offer a fair amount of magneticdipole at low power consumption and mass [47, 48]. The electrical interfacerequires only a driver with a small number of COTS components. Three typesof magnetic coils are commonly found on CubeSat missions:

Wound torque rods are made of a ferromagnetic cylindrical core with enam-eled or self-bonding copper wire wound around. They provide largemagnetic moment per volume and electric power, respectively, butintroduce residual magnetic dipole due to hysteresis of their core.

Wound air coils do not feature a ferromagnetic core, and are either mountedto the satellite structure or installed on solar panels. Their outerdimensions are in most cases prismatic. These actuators produce lessmagnetic moment per volume and electric power. However, they do notemit a residual magnetic dipole.

Embedded air coils are fabricated into the multi-layer PCBs of CubeSatsolar panels. Embedded air coils have properties similar to wound aircoils with regards to magnetic behavior and power consumption.

During design and development of a CubeSat mission, ADCS engineers meetthe task of finding the optimal magnetic actuator. This leads to a comparisonbetween mission requirements and properties of different magnetic actuators,either available on the market or to be developed mission-specific. The fourmain actuator properties for evaluation are magnetic dipole, electric powerconsumption, mass, and volume.

Comparison of magnetic actuators in terms of magnetic dipole and powerconsumption is straightforward. Volume calculation of wound torque rods isuncomplicated and provides a good measure of comparison between different

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38 4 Magnetic Actuator Optimization

wound torque rods due to their compact design. The void area in the centerof wound air coils makes volumetric comparison more complicated. With theheight being the thickness of the PCB substrate, embedded air coils featuresmall volumes but have even smaller void areas in the center of the coil,creating additional difficulties for volumetric comparison with wound air coilsand torque rods.

Problems arise while comparing actuator masses, as masses stated in literatureor by CubeSat component manufacturers sometimes include the mass ofmounting fixtures, sometimes not. Especially for embedded air coils thequestion is asked, if the mass of the solar panel PCB is accounted for in thetotal actuator mass or not. The same then holds for tertiary structure likescrews or nuts that are used to mount the solar panel to the satellite structure.

This chapter aims at providing ADCS engineers with a method to find asolution to the optimization problems encountered during the developmentprocess of magnetic actuators, and also a means to compare different magneticactuator types during mission development.

The content of this chapter is based on material published by the author andSuchantke in [75] and Suchantke in [76]. Those publications are consideredto be in continuation of the work presented by Yoon in [77], which lead tothe realization of the solar panels used on the family of BEESAT satellitesdeveloped at TU Berlin.

4.1 Magnetic Actuator State of the Art

Due to their seemingly simplicity, design, development, and optimization ofmagnetic actuators is subject of many Bachlor’s and Master’s theses. For thesame reason, they are commonly used on CubeSat missions and are offered ina broad variety in CubeSat shops.

To have an overview of the main properties of magnetic actuators, in-depthresearch was conducted. Over forty magnetic actuators designed for or used onCubeSat missions were gathered, and analysis of their properties was presentedby Grau and Suchantke in [75]. The present work extends the existing database

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4.1 Magnetic Actuator State of the Art 39

so that now properties of about one hundred coils are available for analysis.The complete list of coil properties is found in the appendix (cf. appendix A).An updated reproduction of the overview presented in [75] is given in figure 4.1.Displayed are magnetic dipole 𝜇, electric power consumption 𝑃 , and mass𝑀 of the coils, color-coded by coil type. Actuator mass 𝑀 is proportionalto the area enclosed by the circles. In addition to the rectangular main grid,the figure shows a second grid that represents the dipole-to-power ratio 𝜇

𝑃 ,labeled on the top and right border of the diagram.

4.1.1 Wound Torque Rods

All wound torque rods, except for coil number (34), have a dipole-to-powerratio of better than 2:3. Coils (33), (34) and (35) are offered by ZARMTechnik AG and advertised to be optimized for mass and bound by additionalconstraints for minimum magnetic dipole or maximum power consumption.This is a first hint, that careful optimization of coil parameters allows flexibilityin coil design even under rigid constraints. Remaining wound torque rodsappear to be poorly optimized in terms of mass in comparison.

Coil (43) seems to outperform all of the other wound torque rods in terms ofdipole-to-power ratio, which is close to 5:1. However, this coil was designedduring a CubeSat design competition and its properties need to be treatedcarefully [78].

Properties for coil (54) are taken from a report that claims that an optimizationapproach was used during coil design [79]. But like the other coils in thisregion, this one seems to be poorly optimized for mass.

4.1.2 Embedded Air Coils

Embedded air coils are found to have dipole-to-power ratios of less than 1:2(cf. figure 4.1). Coil masses are all less than 10 g, however, masses for coils(49) and (50) are estimated to 6 g as no coil masses were available in themanufacturer data sheets[80, 81]. Both coils offer a good dipole-to-power

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4.1 Magnetic Actuator State of the Art 41

ratio of 1:2, with coil 49 producing less magnetic dipole. Coil 50 operatesfrom a higher power level and therefore generates a higher maximum dipole.

Coils (21) to (24) are all developed for the use on the TU Berlin BEESATmissions, which explains their close resemblance, and coils (23) and (24) haveflight heritage on satellites of the BEESAT family [77, 82].

Coil (28) was designed for the MOVE-II mission of TU München. This missionis planned to be launched in 2018 [83]. In comparison to the previouslydiscussed embedded air coils, those five coils seem to be poorly optimized interms of dipole-to-power ratio, which stays well below 1:4 for all of them.

4.1.3 Wound Air Coils

The region between embedded air coils and wound torque rods in figure 4.1 isfilled by wound air coils. This type of coil varies greatly in terms of dipole-to-power ratio, which is found to be between 1:4 and 2:1. But also their massesshow a large range of variation. Due to the diverse properties, discussionon wound air coils will focus on some highlighted candidates rather thanaddressing all presented coils.

Coil (9), developed for and flown on CanX-1, shows good sense in designingwound air coils already for one of the first CubeSats ever, launched in 2003[84]. The actuator has a good dipole-to-power ratio of about 2:3 and a massof less than 12 g. Coil (10) realizes a similar design for the Belgian CubeSatOUFTI 1 [85].

Best dipole-to-power ratio of all wound air coils is found for coil (19) [86].This coil also shows the largest mass of all coils in the scope of the analysis.The slender and long geometry of the coil, uncommon for wound air coils,requires a large amount of support structure.

Smallest mass of all wound air coils was found for coil (1) which is used asactuator on the UWE-3 satellite [87]. This coil has half of the dipole-to-powerratio of coil (9) at half of the mass. It was designed to be lightweight andgenerate a rather small magnetic dipole in favor for mass and geometricdimensions. On orbit, this design strategy was found to be problematic as due

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to unfortunate circumstances the satellite exhibits a residual magnetic dipolewhich is very close to the one the actuators are able to produce (cf. [88]).

An interesting behavior is seen for coils (5) to (8), which were developed forthe COMPASS-1 mission. The four different entries in the diagram are infact one coil with dipole and power calculated for temperatures of −60 ∘C,0 ∘C, 50 ∘C and 100 ∘C [89]. To better distinguish them from the other coils,they are connected by a thin line. Entry (5) is for the coldest temperature,entry (8) for the hottest. Due to the non-linear dependence of dipole andpower on the coil temperature, the distances between neighboring coils onthat line get closer for increasing temperatures. The dipole-to-power ratiostays constant over the temperature range, which means that coil quality doesnot degenerate.

Coils (16), (17) and (18) are taken from a dissertation researching optimizationof all magnetic actuators [90]. Starting from coil (18), those three coilsapparently show a way how to decrease the power consumption by sacrificingmagnetic dipole. With a dipole-to-power ratio of 1:2 and a mass of 9.90 g, coil(16) would already be a good candidate for 1 U CubeSat missions. However,the optimization of magnetic actuators presented in [90] is highly doubted.While the work presents a well-founded mathematical description of coilproperties (cf. section 4.2.2), the allowed degree of freedom in the parameterspace appears to be overly constrained and no in-depth analysis of the relatedconsequences is provided.

4.2 Magnetorquer Optimization State of the Art

The evidently poor optimization of magnetic coils has already been subject ofresearch in other work (cf. [79, 90, 91, 92, 93, 94]). From experiment-basedoptimization over best-guess to using more advanced methods like sequentialquadratic programming, a multitude of methods is available. This sectionprovides an overview of the state of the art of magnetorquer optimizationavailable in literature and discusses the applicability of each optimizationapproach.

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4.2 Magnetorquer Optimization State of the Art 43

4.2.1 Experiment-Based Optimization

Dildar et al. in [91] and in succession Ali et al. in [92, 93, 94] present the designof reconfigurable embedded air coils for CubeSat applications. Reconfigurablecoils allow to study the impact of multiple sets of magnetic and electricproperties on small satellite missions. With this approach, optimization wouldtake place in the laboratory or on orbit by experimenting with different setupsfor those configurable coils.

Experiment-based optimization is inefficient, as appropriate coil candidatesneed to be identified using best guesses and manufactured before they can betested in the laboratory. So other methods that depend on the mathematicalmodeling of magnetic actuators are superior to experimental methods.

4.2.2 Model-Driven Optimization

Model-driven optimization in general is based on mathematical models thatdescribe the system to be optimized. They are then used together with anoptimization technique to find optimum candidates. Model-driven optimizationof magnetic actuators is applied by Bellini in [90]. His work starts with thedefinition of mathematical models of embedded air coils, wound air coils, andwound torque rods. From each coil type, a single specimen is manufacturedand subjected to test, and the mathematical models are validated against theresults. Based on the mathematical models, a design tool is implemented,which promises to find an optimal candidate from all three coil types. Thetool uses a three-staged approach to finding the optimal candidate, whichalso incorporates consideration of manufacturing costs.

The work presented by Bellini in [90] is criticized for multiple aspects. First,his mathematical models use many simplifications instead of implementing amore realistic model. This leads to the introduction of corrective factors tocompensate the inaccuracies.

Second, the model for calculating embedded air coil magnetic dipole is appliedto all layers in the PCB, which would lead to a coil that does not create anymagnetic dipole, due to alternating winding sense of the current through the

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coil. Bellini further identifies a residual magnetic force for a single layer, andproposes to use the same layout rotated by 90° in four subsequent layers. Thiswould render coils with a number of layers not divisible by four infeasible.

Third, only a small number of variable input parameters are taken into accountfor parameter optimization in the presented use cases [90]. This leads to anover-constrained parameter space, and prevents better optimization.

Finally, properties of the coils resulting from the optimization procedure appearto be poorly optimized in comparison to the coils shown in figure 4.1. Dipole-to-power ratio of all resulting coils is below 1:2, and masses are higher thanthe average in the respective class. Of the optimized embedded air coils, nocoil is even shown inside the diagram area.

4.2.3 Sequential Quadratic Programming

Optimization of magnetic actuators using a sequential quadratic programming(SQP) method to find the maximum magnetic dipole for a given powerconsumption and mass of a wound torque rod is presented by Miller in [79].Those methods require the formulation of a cost function, and find theminimum of this function from an initial guess of parameters, which are boundby a set of additional equality and inequality constraints. In [79], the costfunction expresses magnetic dipole. Four inequality constraints are used toconstrain mass, power, total number of turns, and ratio of core length toradius. As pointed out in the original source, the applied optimization methodhas several limitations which are summarized here:

– SQP may only be used with continuous parameters.

– Depending on the initial guess, only a local optimum may be found.

– For infeasible problems, the SQP method attempts to minimize themaximum constraints, leading to a violation of the constraints.

Some samples of optimized coils are presented in [79]. Well optimized coilsfrom there have good dipole-to-power ratios of above 1:1, but feature ingeneral larger masses as the coils presented in figure 4.1.

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4.3 Formulation of a Novel Optimization Procedure 45

Especially the inability to use discrete values for input parameters to thecost function is criticized. For a real world magnetic actuator the enameledwire will only be available in discrete diameters. And for embedded air coilsthe number of turns per layer and the number of layers are also discreteinteger values. Finding optimized solutions for such a setup requires the useof so-called mixed-discrete sequential quadratic programming methods. Theyare more difficult to apply and have even bigger problems in finding the globaloptimum.

A solution to the problem of finding the global optimum is not addressed byMiller in [79], and no proof is shown, that the selected coil actually representsthe global minimum. One step towards that direction is the discussion led onthe complex shape of the design parameter space, where the proposed costfunction was found not to be strictly convex.

4.3 Formulation of a Novel Optimization Procedure

From the discussion in the previous section, the following requirements for anovel optimization procedure are derived for this work:

– Precise mathematical models of magnetic coil properties shall be usedwith the optimization procedure. Models that are based on estimatesor mean values shall be avoided, as they may lead to misconceptionsduring preliminary design phases.

– The procedure shall use as many design parameters as possible tohave a high-dimensional parameter space. With modern day computersand engineering programming languages, a computationally intenseprocedure is no problem to be realized.

– It shall enable the use of discrete input parameters like wire gauges,copper weights, or supply voltages. This better represents the availabletechnologies for coil manufacturing and eliminates the process of rehash-ing the identified parameters of an optimal solution for manufacturing.

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– A flexible interface for defining optimization objectives shall be imple-mented by the procedure. Equality and inequality constraints on theinput as well as on the output parameters and the possibility to searchfor minimum and maximum values of the objectives need to be set usingthis interface.

– Owing to the high dimensionality and the resulting complex shape of theparameter space, the procedure shall be able to find the global optimumfor a given optimization scenario.

How the above requirements are met by the proposed optimization procedureis shown in the following subsections. Research related to those sections waspublished by Grau and Suchantke [75], which relies on the work presented bySuchantke in [76]. For the present work, the previously published work wasreviewed and the implementation further improved.

4.3.1 Mathematical Modeling of Magnetic Coil Properties

To meet the first requirement, a complete as possible set of input designparameters was identified for all three coil types. Input design parametersare e.g. wire diameter 𝑑w, core length 𝑙c and diameter 𝑑c, and parametersthat describe the geometry of the coil, but also material coefficients like thetemperature coefficient of specific resistance 𝛼 or density 𝜌. Those werethen used to define equations that put the output design parameters intorelation with the input parameters. The main output design parameters aremagnetic dipole 𝜇, electric power consumption 𝑃 , and mass 𝑀 . Whereverpossible, basic input parameters are first linked together to form intermediateparameters, like e.g. wire length 𝑙w or enclosed area 𝐴 of the coil. The reasonbehind defining intermediate parameters is, that some common relations likeOhm’s law to calculate the current 𝐼 from supply voltage 𝑈 and resistance 𝑅

are the same for no matter what type of coil.

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4.3.2 High-Dimensional Parameter Space Implementation

The second requirement demands a large number of input parameters in orderto generate a high-dimensional parameter space. Owing to the many discreteor integer-valued input parameters and the additional implementation of aglobal minimum search method, a mixed-integer SQP method was rejected tobe used for the optimization procedure. Instead, a brute force approach wasselected, which exploits the matrix calculation capabilities of the engineeringprogramming language GNU Octave [95].

The equations previously discussed are all implemented as functions that allowparameters to be passed as multi-dimensional vectors into the function. Asthe input parameter vectors are originally defined to be one-dimensional, theyare all reshaped to matrices of a fixed number of dimensions, with each matrixhaving only multiple entries in one single dimension. The functions then useelement-by-element operators and broadcasting mechanisms of GNU Octaveto calculate the output as multi-dimensional matrices.

The advantages of using this approach over coding in a high-level, object-oriented programming language like e.g. C++ are the ability to quicklyimplement new features during development due to the unnecessary compileand link steps. While GNU Octave has the great advantage of being able towrite program code that is very similar to the mathematical formulation ofthe problems, it also supports test-driven development. And finally, it givesaccess to a great reservoir of high-level plotting functionality. Most of thediagrams displayed in this work were in fact created in GNU Octave and runthrough a self-developed tool that translates the diagrams into high-qualityvector graphics.

One of the disadvantages of using GNU Octave is, that it is a scriptinglanguage and an interpreter is used to execute the code. In comparison to acompiled and linked program written in a high-level programming languagelike C++, the code is therefore executed much slower. However, as GNUOctave itself is written in C++ and the functions developed in the scopeof the optimization procedure make heavy use of the element-by-elementoperators, which are directly implemented in C++, those speed disadvantagesare attenuated.

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In order to be able to calculate intermediate and output coil parameters forthree coil types and to display the data in diagrams and tables, over onehundred functions were implemented. They are structured into functions usedfor individual coil types, as well as common functions like e.g. reading wiredata or display intermediate and output parameters in tables and diagrams.

4.3.3 Discrete and Integer Input Parameters

In order to fulfill the third requirement, input parameters are defined in GNUOctave script files. The parameter file is passed to a high-level function whichcalculates the intermediate and output parameters and takes care of plottingand generation of tables.

Listing 4.1 gives an example for such an input parameter definition file, takenfrom the optimization of an embedded air coil. All input parameters are givenin SI basic units, like e.g. m, V, or ∘C. Ranges of input parameters are eitherdefined as GNU Octave ranges, which require the base, increment, and limitto be set, or row vectors that define a set of integer or decimal values. Someparameters, like e.g. embedded air coil board length 𝑎 are fixed to a singlevalue, but in order to reduce the amount of typing necessary to quickly adaptthe setup, the structures to define ranges are kept.

This allows to gain very granular control over the true dimensionality andlevel of detail of the parameter space. It also enables to represent certaintechnological aspects. Take e.g. the definition of supply voltage 𝑈 : The entries1.80 V, 3.30 V and 5 V represent common microcontroller supply voltages forwhich voltage converters might already be implemented in hardware. Batteryvoltage levels then relate to the entries 3.70 V and 7.40 V. Thus, the parameterdefinition in listing 4.1 allows to identify which supply voltage out of a set ofavailable ones would be optimal.

4.3.4 Flexible Optimization Objectives Definition

In order to quickly address the introduction of new optimization objectivesand to define if a parameter should be minimized or maximized, optimization

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Listing 4.1: Example of discrete and integer input parameter definition in GNUOctave

% coil dimensionsa = ( 94: 2: 94) * 1e -3;b = ( 80: 2: 80) * 1e -3;d = 200e -6;

% PCB manufacture parameterss = (150: 25: 150) * 1e -6;w = (150: 10: 260) * 1e -6;h = [18 35 70 105 140] * 1e -6;

% coil designk = ( 40: 1: 40);m = ( 4: 1: 4);

% supply voltageU = [1.8 3.3 3.7 5 7.4];

% temperature ranget = ( 20: 5: 20);

objectives are defined as cell matrices. An example of such a definition takenfrom the optimization of an embedded air coil is shown in listing 4.2.

The first three entries in the objective definition relate to finding optimalsolutions for magnetic dipole 𝜇, power consumption 𝑃 , and mass 𝑀 . Magneticdipole is to be maximized, starting from a minimum value of 27.5 mA m2.Power and mass are to be minimized, starting from maximum values 83.3 mWand 5.83 g, respectively. Above objective definition finds a total of fiveoptimized solutions. In case of the magnetic dipole optimized coil, the otherfour objectives act as inequality constraints. So in this case, the coil will havethe most magnetic dipole of all calculated coils while obeying the upper limitsfor power, mass, and so on.

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Listing 4.2: Example of flexible optimization objective definition in GNU Octaveopt = {

’mu ’, ’max ’, 27.5e-3’P’, ’min ’, 500e-3 / 6’M’, ’min ’, 35e-3 / 6’m’, ’min ’, 7’Ai ’, ’max ’, 1e -12

};

The fourth entry states the inequality constraint 𝑚 ≤ 7, which means that allcoils occupy no more than seven internal layers in the PCB. It is possible todefine the same inequality constraint using only the input parameter rangedefinitions as given in section 4.3.3. Adding a fourth definition to the list ofoptimization objectives, however, will result in an optimal solution that obeysall other inequality constraints, and has the smallest possible number of layers.As the number of internal layers directly influences the cost of the coil, thismight be a way to also find a cost-optimal solution. A similar considerationwas discussed by Bellini in [90], but no solution to the problem provided.

The last entry in listing 4.2 is important for highly integrated CubeSats solarpanels, which feature a large number of electric components. To enablesetting this objective, an additional equation for the area 𝐴𝑖 enclosed by theinnermost copper traces of the coil was derived and implemented in GNUOctave. In order to now maximize the available inner area, this objective isadded which will result in a fifth optimal solution to the problem. To not biasthe results for the first four objectives defined, the minimum value for theinner area is chosen at a very small value.

4.3.5 Global Optimum Search

Following the calculation of output parameters of a large set of coil candidates,which are stored in multi-dimensional matrices, optimal solutions need tobe identified. The implemented search uses GNU Octave’s find functiontogether with a combination of boolean and comparison operators to identify

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the optimal solutions. The search string is automatically created from thedefined optimization objectives. An example of such an automatically createdsearch string that uses the optimization objectives defined in listing 4.2 isgiven in listing 4.3.

The comparison expressions ≥ and ≤ are applied to the output parametermatrices individually. In a previous step, all parameter matrices have beenbroadcasted to the same size. Therefore, return values of all comparisonexpressions have the same size as well. Boolean expressions are applied next,which yields a single matrix of predefined size that holds ones in all entriesthat obey the constraints. The find function then returns a single row vectorof indexes into the parameter matrices.

Next, for each objective the relevant maximum or minimum is located in theoutput parameter matrices. This is spelled out in listing 4.3 on the last fivelines, but automated in the source code, to maintain flexibility in terms ofoptimization objectives definition. The variable iopt finally holds all indexesthat represent optimal candidates which obey the inequality constraints.

As the implemented procedure does not use a continuous parameter space, itis not able to locate the mathematically exact global optimum. But it willidentify the discretized optimum solution closest to the exact global optimum.

Listing 4.3: Example of global optimum search in GNU Octaveind = find ( (mu >= 0.0275) ...

& (P <= 0.08333) ...& (M <= 0.005833) ...& (m <= 7) ...& (Ai >= 1e -12));

[~, iopt (1)] = max (mu(ind ));[~, iopt (2)] = min (P (ind ));[~, iopt (3)] = min (M (ind ));[~, iopt (4)] = min (m (ind ));[~, iopt (5)] = max (Ai(ind ));

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As one of the basic requirements in the design of the procedure was to be asclose as possible to a real-world design problem for magnetic coils, this is notconsidered to be a drawback. Like with the previously discussed optimizationprocedures, the user still has to set up the input parameters carefully in orderto not constrain the parameter space too much.

4.4 Optimization Results

Magnetic actuator optimization results based on the procedure derived insection 4.3 were first published by Grau and Suchantke in [75]. All resultsavailable there have been added to the existing magnetic actuator database,and an updated diagram is shown in figure 4.2.

4.4.1 Wound Torque Rod Optimization Results

A first attempt for coil optimization was made based off of coil (34) [96]. Thisresulted in coils (a), (b), and (c), where the parameters of coils (b) and (c)are identical and therefore only coil (b) is displayed [75]. The magnetic dipoleoptimized variant (a) offers a small gain in terms of magnetic dipole and alsopower consumption. Power optimized variant (b) reduces power consumptionby about 15 %.

The results for wound torque rods need to be treated very carefully, as theyare only simulated and the magnetic behavior of the core might introduce alarger deviation between modeled and measured magnetic properties.

4.4.2 Wound Air Coil Optimization Results

Optimization of a wound air coil was based off of coil (1) [87, 75]. The threeresulting coils (d), (e), and (f) are very close to the targeted coil. Poweroptimized variant (e) shows the biggest potential for optimization with areduction of required power of about 13.6 %.

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The confidence in the optimization results depends on the experience withthe manufacturing process. If the winding process of rectangular coils isunderstood and applied correctly, it is possible to achieve the optimizedgeometry closely.

4.4.3 Embedded Air Coil Optimization Results

Constraints in [75] for embedded air coil optimization are defined based oncoils 23 and 24 [82]. Obeying those constraints, embedded air coils show bigpotential for optimization. The dipole optimized solution (g) is very close tocoil (23), but with a mass reduction of about 2 g. A dipole-to-power ratioof greater than 1:3 is achieved with coil (h) which gets close to coil (49).Mass optimized coil (i) shows, that masses of down to 1.84 g are possible,which is equivalent to a reduction of 69.3 %. The most interesting candidateis coil (k), hidden beneath coils (i) and (j), as this coil only requires threeinternal layers. This would enable the design of a solar panel that requires onlyfour or six layers, which in turn will reduce manufacturing cost drastically incomparison to the ten-layer PCBs that are currently used for the solar panelsof the BEESAT satellites.

Application of the optimization procedure shows, that embedded air coils offerbig potential for optimization. There is not only potential to increase themagnetic dipole or decrease power consumption or mass, but also to reducemanufacturing costs by finding solutions that allow to be implemented on lesslayers. Confidence in the optimization of this coil type is very good, as thechemical etching of copper traces is well reproducible and the vacuum coredoes not introduce any uncertainties. This also shows, that embedded aircoils have a huge potential for developing custom coils for individual missionsfrom scratch given a set of constraints and technology parameters.

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5 A New Class of Picosatellite AttitudeActuators

Unlike reaction wheels, which store angular momentum using rotating flywheels,fluid spacecraft actuators employ a fluid which is pumped through a closed-loop conduit. First fluid actuators were proposed in the late 1950’s. Butonly recently, implementation of so-called fluid-dynamic actuators on pico andnano satellite missions came into scope.

5.1 Technological Evolution of Fluid Spacecraft Actuators

Haviland in his 1958 patent [97] claimed a device for stabilizing the attitudeof a spacecraft employing a dense liquid pumped in an endless conduit. Beingable to propel the liquid in a reversible and controlled manner allows tocounteract external disturbances on the spacecraft and therefore to controlits attitude. The patent presents an embodiment of an electromagnetic pumpconsisting of a magnetic field and a pair of electrodes which fit into a conduitof circular cross section. The generated electric motor effect causes a forceon the electrically conducting liquid and accelerates it.

A fluid mass gyroscope, consisting of a fluid ring actuator and single or doublegimbals, was claimed by Yeadon in his 1960 patent [98]. A double-gimbaledfluid gyroscope allows to generate control torque about two axes for a fixedvelocity of the fluid.

Maynard in his 1988 patent [99] claimed a fluid momentum controller toneutralize external torques exerted on spacecraft and other vehicles. Hisinnovation uses a small mass of pumped fluid in small-diameter pipes, thatare located at the periphery of a vehicle’s structure to generate large controltorques. Maynard also claimed an electromagnetic pump of prismatic shape.

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56 5 A New Class of Picosatellite Attitude Actuators

He derived, that for a planar fluid momentum controller of rectangular shapeangular momentum storage capability is proportional to the enclosed area ofthe loop, the fluid velocity, fluid density, and the pipe cross-sectional area.

Iskenderian in 1989 [100] and Lurie and Schier in 1990 [101] extended researchfluid-loop reaction systems. Lurie, Schier, and Iskenderian [102] in 1991claimed the invention of multiple fluid paths of either circular or irregularshape that are located on the interior or exterior of a spacecraft to realize fullattitude control along any given axis. Using different valves, they describea tetrahedral fluid path network. Switching fluid between triangles allows togenerate three different torque vectors with less investment in piping.

The aforementioned studies and patents only describe concepts and do notinvolve feasibility analyses for real applications or hardware demonstration.A first experimental investigation was conducted by Kelly et al. [103] in2004. The experiment was conducted in the scope of NASA’s reduced gravitystudent flight program. Their prototype featured two fluid loops with parallelaxes of symmetry, and therefore was only able to produce torque about asingle axis.

In 2009, Kumar in [104] researched three-dimensional spacecraft attitudestabilization using three fluid ring actuators. He derived linearized systemmodels and developed linear and nonlinear control laws based on sliding modecontrol techniques. Numerical simulation proved the feasibility of achievingspacecraft attitude control with three fluid rings, even in the presence of highdisturbance torques and intermittent actuator failures.

In her 2013 doctoral thesis [105], Nobari stated that "despite its significance ...a detailed analysis of the feasibility of using fluid rings as attitude actuators aswell as their failure analysis has not been reported in the literature". Investigat-ing the dampening effect of fluid rings on spacecraft attitude numerically, shefound that passive stabilization alone is not enough to detumble a spacecraftin low earth orbit in reasonable time. She then compared the use of a PID anda sliding mode law to control the three-dimensional attitude of a spacecraft.To show the advantage of using a redundant set of fluid rings, a failure analysiswas performed. As sliding mode controllers tend to chattering in the steadysystem response, the last part of her analysis included the design of a switchingcontroller combining a sliding mode with a PID control law.

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5.2 Objectives for Actuator Miniaturization 57

The experimental part of Nobari’s work covers the conduction of single-axisas well as three-axis experiments on a floating air simulator. Results fromthe single fluid ring system proved the concept of using such actuators. Forthe three-axis setup, Nobari concluded that the applications of fluid rings asauxiliary actuators is limited due to high required voltages and small resultingangular rates of the fluid.

One solution to the problem of small fluid angular rates is to use a metallicfluid of high density. In his 2011 patent [106], Noack claims multiple differentembodiments of what he calls fluid-dynamic actuators. The differences betweenhis invention and earlier patents are mainly in fluid used to store angularmomentum, pump and related electronics, as well as shapes of fluid conduits.Noack in [107] describes the development and test results of a fluid-dynamicactuator to be used on the TechnoSat mission of TU Berlin. In [108], Noackshowed the possibility to miniaturize fluid-dynamic actuators (FDAs) for theuse on pico satellite missions with a laboratory demonstrator that uses aconduit fit for CubeSat application.

5.2 Objectives for Actuator Miniaturization

Small fluid spacecraft actuators are considered to provide a solution to thechallenges encountered using miniaturized reaction wheels. So-called fluid-dynamic actuators are based on high-density metallic fluids and do not showthe disadvantages of reaction wheels due to the fluid being the only movingcomponent in the actuator. There is no need for ball bearings, no wear, andtherefore no lubricant. In addition, friction is dictated by shear stress in theliquid, which results in continuous provision of torque at small angular ratesand does not introduce any kind of jitter due to imbalance. Figure 5.1 showsan image of the first laboratory demonstrator presented by Noack in [108] onthe air bearing test facility at TU Berlin.

The fluid conduit of this actuator is made of PVC tube and the electroniccomponents are soldered to a piece of veroboard. A simple proportionalcontroller is implemented on an 8-bit microcontroller, and allows the studyof attitude control maneuvers on the air bearing platform. Noack showed in

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Figure 5.1: First demonstrator of a CubeSat fluid-dynamic actuator on an air bearingtest facility [108]

[108], that fluid-dynamic actuators produce very high torques in comparisonto reaction wheels, but are restricted in terms of angular momentum storagecapacity.

Due to its nature of being a functional laboratory demonstrator, the actuatoris not designed to be used on CubeSat missions. Its toroidal shape interfereswith the mostly cubic envelopes of CubeSat subsystems and components.This leads to poor overall volumetric utilization on the spacecraft, thereforelimiting available payload volume.

Improving the volumetric utilization and integration of fluid-dynamic actuatorswith CubeSats requires the identification of a suitable conduit geometry (cf.section 5.4). To adapt the demonstrator for a better fit with a cuboid envelope,manufacturing technologies for fluid-dynamic actuators in [107] are rejecteddue to the stricter constraints found on CubeSats in terms of volume and mass.So development of a conduit geometry suitable for CubeSat applications ispaired with the selection of an appropriate, modern manufacturing technologythat provides maximum design flexibility.

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5.3 Fluid-Dynamic Actuator Fundamentals 59

Miniaturization of the picosatellite fluid-dynamic actuator electronic compo-nents (cf. section 5.5) is following identification of a suitable conduit geometryand manufacturing process. Every actuator is equipped with a microcontroller,that interfaces with the other subsystems in the satellite over a common busprotocol. This allows to close the control loop using additional sensors andactuators as well as multiple control schemes in the ADCS. Please note, that itis not the goal of this thesis, to develop high-level embedded control softwarefor the picosatellite fluid-dynamic actuator electronics, but to implement thebasic functionality in software and enable access to the low-level functionalityof the actuator electronics for following functional verification over the businterface.

Functional verification of the developed picosatellite fluid-dynamic actuatorincludes the verification of the dynamical properties of the actuator on an airbearing, as well as the measurement of the power demand (cf. section 5.6).Utilization of a low-cost air bearing introduces disadvantages like perturbationtorque induced by the air flow in the bearing. Measured data is separated fromperturbations in order to estimate the dynamical properties of the actuators.

Redundancy concepts based on FDAs have not been addressed in the literatureso far. Modern manufacturing processes allow for a great flexibility in conduitgeometry, enabling possibilities for redundancy using tetrahedron or pyramidalassemblies of pFDAs (cf. section 5.8).

5.3 Fluid-Dynamic Actuator Fundamentals

Maynard states (cf. [99], section 5.1), that for a planar fluid momentumcontroller of rectangular shape angular momentum storage capability is pro-portional to the enclosed area of the loop, the fluid velocity, fluid density, andthe pipe cross-sectional area. In general, the angular momentum vector 𝐻 ofa point mass 𝑀 is given as cross-product of radius vector 𝑟 and momentumvector 𝑝, where the momentum vector is given as product of mass timesvelocity vector 𝑣 :

𝐻 = 𝑟 × 𝑝 = 𝑀 𝑟 × 𝑣 (5.1)

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An alternative formulation for rigid bodies uses the mass moment of inertia 𝐼

and the angular rate vector 𝜔 :

𝐻 = 𝐼 𝜔 (5.2)

For rotational motion, every point mass that constitutes a rigid body istraveling at the same angular rate. For fluids in a conduit, this only holds trueif the conduit guides the fluid on a circular path, and no shear movement takesplace. For any other conduit geometry, above equation is not applicable, andthe angular momentum needs to be approximated using a sum over multiplefluid elements, that constitute the conduit geometry:

𝐻 =∑︁

𝑖

𝑀 𝑖𝑟𝑖 × 𝑣 𝑖 (5.3)

Here, 𝑀 𝑖 is the mass of each element, 𝑟𝑖 is the position of the center of massin three-dimensional space, and 𝑣 𝑖 is the velocity vector, approximated to beuniform over the volume of each element.

Under the assumption, that cross-sectional dimensions of a conduit are verysmall with respect to its overall geometry, calculation of the angular momentumvector is greatly simplified. Now, linear elements of the conduit geometrycan be treated separately, and then be summed up to calculate the angularmomentum according to eq. (5.3).

5.3.1 Planar Actuators

Planar actuators have the conduit structure and therefore the fluid moving ina plane. For simplicity, the xy plane is considered to be the reference planein which the conduit is located, and the z axis is then perpendicular to theconduit structure. Hence, the angular momentum vector points in positive zdirection, if the fluid moves in a counterclockwise direction in the xy plane.

Planar actuators allow for simple mechanical interfacing between the actuatorsand the CubeSat structure. Actuators might be placed in the PCB stack, ormounted on solar panels (cf. chapter 6). Three-axis attitude control is thenachieved using three planar actuators, with their axes of angular momentumpointing along the three spatial directions.

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5.3.1.1 Circular Conduit

For a circular conduit as shown in figure 5.2, the distance 𝑟 of a mass point𝑑𝑀 on the circle is always the same with respect to the center, and thevelocity 𝑣 is always perpendicular to the radius. This allows to write the totalangular momentum of the fluid in integral form using scalars:

𝐻 =∫︁

𝑀

𝑟𝑣 𝑑𝑀. (5.4)

Using the density 𝜌 of the fluid, the cross-sectional area of the conduit 𝑆, andthe small angle 𝑑𝛼, eq. (5.4) is rewritten as

𝐻 = 𝑣𝜌𝑆𝑟2∫︁ 2𝜋

0𝑑𝛼. (5.5)

This results in

𝐻 = 2𝑣𝜌𝑆𝜋𝑟2. (5.6)

Using the enclosed area 𝐴 = 𝜋𝑟2 yields

𝐻 = 2𝑣𝜌𝑆𝐴. (5.7)ii

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𝑑𝑀

𝐴

𝑟

𝑣

Figure 5.2: Circular conduit

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Equation (5.7) proves that the assumption of Maynard (cf. section 5.1) forrectangular conduits holds also for circular conduits.

Assuming uniform fluid motion about the z axis, angular momentum is calcu-lated from mass moment of inertia 𝐼 and angular rate 𝜔. Moment of inertiaof the circular conduit is given as

𝐼 = 𝑀𝑟2. (5.8)

Rewriting above equation using fluid density, cross-sectional area, and lengthof the streamline yields

𝐼 = 2𝜌𝑆𝜋𝑟3. (5.9)

For a circular conduit, angular rate is defined by fluid velocity and radius as

𝜔 = 𝑣

𝑟. (5.10)

Combining above two equations results in

𝐻 = 2𝑣𝜌𝑆𝜋𝑟2. (5.11)

This proves, that for circular conduits both ways to estimate the angularmomentum are identical.

5.3.1.2 Rectangular Conduit

Rectangular conduits allow to optimize the volumetric utilization of pFDAs byadapting to the cubic shapes of adjacent subsystems. For a rectangular shapewith width 𝑎 and height 𝑏 as shown in figure 5.3, angular momentum is givenin [99] as

𝐻 = 2𝑣𝜌𝑆𝑎𝑏 = 2𝑣𝜌𝑆𝐴. (5.12)

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𝑑𝑀𝐴

𝑟

𝑣

Figure 5.3: Rectangular conduit

5.3.1.3 Square Conduit

Angular momentum of square-shaped conduits is derived from the rectangularshape by setting 𝑎 = 𝑏:

𝐻 = 2𝑣𝜌𝑆𝑎2. (5.13)

5.3.2 Three-Dimensional Actuators

Three-dimensional shapes have the conduit and therefore the flowing liquidnot only in a single plane. This allows to guide the fluid along the edges of theCubeSat outer structure. L-shaped conduits have conduits in the xy and yzplanes, and crown-shaped conduits have conduits in the xy, yz, and zx planes.Angular momentum vectors of these conduit shapes are no longer parallel tothe three spatial directions. Proper dimensioning of the conduit geometryenables redundant actuator designs similar to tetrahedron configurations ofreaction wheels with an increased volumetric utilization (cf. section 5.8).

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𝑥

𝑦

𝑧

𝐻

𝐻𝑎𝑐

𝐻𝑏𝑐

𝑐 𝑎

𝑏Figure 5.4: L-shaped conduit

5.3.2.1 L-Shaped Conduit

Angular momentum calculation for L-shaped conduits as shown in figure 5.4is based on eq. (5.3). Angular momentum vector has each a component inx and z direction. Z component 𝐻𝑎𝑐 is created by the conduit in the xyplane, which spans a rectangle with sides 𝑎 and 𝑐, and x component 𝐻𝑏𝑐 iscreated by the conduit in the yz plane, which in turn spans a rectangle withsides 𝑏 and 𝑐. Angular momentum calculation is simplified, considering therectangular shapes spanned in the xy and yz planes independently, and puttingthe reference point at the coordinate center.

This leads to the angular momentum vector components

𝐻𝑎𝑐 = 2𝑣𝜌𝑆𝑎𝑐, (5.14)𝐻𝑏𝑐 = 2𝑣𝜌𝑆𝑏𝑐. (5.15)

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Angular momentum vector is then given by

𝐻 = 2𝑣𝜌𝑆𝑐

⎡⎣𝑏

0𝑎

⎤⎦ . (5.16)

Magnitude of angular momentum is derived from the vector components as

𝐻 = 2𝑣𝜌𝑆𝑐√︀

𝑎2 + 𝑏2 = 2𝑣𝜌𝑆𝐴 (5.17)

It can be seen from the above equation, that the virtually enclosed area of theactuator is spanned by the elements parallel to the y axis and lines connectingtheir ends. It can be shown, that the angular momentum vector is at all timesperpendicular to that area.

Angle 𝛽 between xy plane and angular momentum vector is calculated fromthe components using

𝛽 = tan−1(︂

𝐻𝑎𝑐

𝐻𝑏𝑐

)︂= tan−1

(︁𝑎

𝑏

)︁. (5.18)

Above equation indicates, that the edge length ratio 𝑎𝑏 can be used to directly

set the angular momentum direction required for redundant configurations (cf.section 5.8)

5.3.2.2 Crown-Shaped Actuators

Crown-shaped actuators, like the one shown in figure 5.5, feature conduitsin the xy, yz, and zx plane. Angular momentum of this conduit geometry isderived in a similar manner as in the previous section. The angular momentumvector features components in all three spatial directions. Z component 𝐻𝑎𝑐

is created by the conduit in the xy plane, which spans a rectangle with sides 𝑎

and 𝑐, and x component 𝐻𝑏𝑐 is created by the conduit in the yz plane, whichin turn spans a rectangle with sides 𝑏 and 𝑐. Last, y component 𝐻𝑎𝑏 is createdby the conduit in the zx plane, spanning a rectangle with sides 𝑎 and 𝑏.

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𝑥

𝑦

𝑧

𝐻

𝐻𝑏𝑐

𝐻𝑎𝑏

𝐻𝑐𝑎

𝑐 𝑎

𝑏Figure 5.5: Crown-shaped conduit

This leads to the components being

𝐻𝑎𝑐 = 𝑣𝜌𝑆𝑐𝑎, (5.19)𝐻𝑏𝑐 = 𝑣𝜌𝑆𝑏𝑐, (5.20)𝐻𝑎𝑏 = 𝑣𝜌𝑆𝑎𝑏. (5.21)

Angular momentum vector is then given by

𝐻 = 𝑣𝜌𝑆

⎡⎣𝑏𝑐

𝑎𝑏

𝑐𝑎

⎤⎦ . (5.22)

Angle 𝛽 between xy plane and angular momentum vector is calculated fromvector components using

𝛽 = tan−1(︂

𝑐𝑎√𝑏2𝑐2 + 𝑎2𝑏2

)︂= tan−1

(︃𝐻𝑎𝑐√︀

𝐻2𝑏𝑐 + 𝐻2

𝑎𝑏+

)︃. (5.23)

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5.3.3 General Spacecraft Dynamics

The following dynamics knowledge is necessary for understanding the procedureand results for pFDA functional verification on the gas bearing. Rigid bodyangular momentum vector in general is given in eq. (5.2). Torque vector 𝜏 isthe time derivative of the angular momentum vector

𝑑

𝑑𝑡𝐻 = 𝜏 . (5.24)

For a rotating rigid body with no external torques, conservation of angularmomentum applies:

𝑑

𝑑𝑡𝐻 = 0, (5.25)

which is expressing, that neither external control torque nor external perturba-tion torque is acting on the rigid body.

For a rigid spacecraft equipped with actuators based on momentum exchange,the dynamic equation is given by

�̇�𝑠 = 𝐼s−1 [𝜏 + 𝜏 𝑑 − 𝜔𝑠 × (𝐼s × 𝜔𝑠 + 𝐻 )] . (5.26)

Here, �̇�𝑠 is spacecraft angular acceleration vector, 𝐼s is spacecraft massmoment of inertia, 𝜏 is actuator torque vector, 𝜏 𝑑 is the disturbance torquevector acting on the spacecraft, 𝜔𝑠 is spacecraft angular rate vector, and 𝐻

is actuator angular momentum vector.

Using a set of three pFDAs, control torque 𝜏 a is given by

𝜏 pfda = 𝑑

𝑑𝑡𝐻 pfda, (5.27)

where angular momentum vector 𝐻 pfda is in turn given by

𝐻 pfda =

⎡⎣𝐻pfda,1𝐻pfda,2𝐻pfda,3

⎤⎦ (5.28)

and 𝐻pfda,1 through 𝐻pfda,1 are angular momentum of the three perpendicularactuators.

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To describe redundant actuator configurations, where the axes of the actuatorsare no longer linearly independent, a distribution matrix 𝑇 is required:

𝐻 pfda = 𝑇 ·

⎡⎢⎢⎣𝐻pfda,1𝐻pfda,2𝐻pfda,3𝐻pfda,4

⎤⎥⎥⎦ . (5.29)

For four actuators, 𝑇 is a 3 × 4 matrix that describes the alignment of theactuators, which is usually tetrahedral or pyramidal.

5.4 Fluid Actuator Conduits for CubeSat Applications

Design of pFDA conduits starts out from the conduit presented by Noack in[108]. Due to its nature, the conduit of Noack’s laboratory demonstrator isnot fit to be directly used on CubeSats.

5.4.1 Conduit Considerations

A circular conduit was designed by Noack against the requirements of havinga maximum diameter of 80 mm, a maximum fluid mass of 20 g, and beingable to store at least 150 µN m s of angular momentum [108]. Application ofeq. (5.7) leads to a required mass flow rate of at least 15 g/s.

pFDA conduit geometry, that better fits the cubic shapes of CubeSat subsys-tems and structures, has square outer dimensions and square cross section.Prospected mounting of three pFDAs on CubeSat solar panels calls for squareouter dimensions with a maximum length of 80 mm to leave enough spacefor CubeSat rails and satellite structure. With fluid mass being limited to20 g and taking the conduit wall thickness into account, conduit cross sectionwidth is calculated. This width is found to be 13 % smaller than the diameterof the laboratory demonstrator conduit. Applying eq. (5.7) again under therequirement for 150 µN m s of angular momentum yields a required mass flowrate of at least 13.4 g/s. This is a reduction of about 10.6 % in comparisonto the conduit presented by Noack in [108].

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5.4 Fluid Actuator Conduits for CubeSat Applications 69

A reduction in required mass flow rate should lead to a reduction in powerdemand of the pump. This is contrasted by the increased friction losses dueto the increased length of the channel and the sharp bends introduced by thesquare conduit. To mitigate the influence of the sharp bends, a square withrounded corners is used as outer shape of the conduit. The resulting conduitgeometry was first presented by Grau et. al in [63].

5.4.2 Pump Housing

Inside the pump, conduit cross-section needs to change from square to flatrectangular (cf. [107]). The large magnetic flux needs to be shielded and istherefore guided using two iron pole pieces. Both, housing and pole pieces,require openings to feed the electrodes into the pump. Electrodes and polepieces need to be electrically isolated from each other.

Design of pump inlet and outlet geometry is intricate: the circular or squareconduit cross-section needs to be transformed to the flat rectangular shapewhile maintaining a constant cross-sectional area. To produce the pump,inlet, and outlet geometry using traditional manufacturing processes likemilling, the pump housing needs to be assembled from two pieces. The twopieces need to have additional features for mechanically aligning and fasteningthem. Additionally, this will result in redundant housing material and thereforeincreased actuator mass. If the conduit is made from tube, additional contactsurfaces are required at the inlet and outlet to the pump in order to glue thetube to the pump housing, which further increases actuator mass.

As a solution to this, the author proposed the monolithic integration of thepump housing into the conduit geometry in [63]. In this design, the pumpsits in one of the corners of the square/square conduit, and free-form inletand outlet directly connect the pump with the channel. Features to align themagnets, feed the electrodes to the pump, align the pole pieces, and electricallyisolate the pole pieces from the electrodes are also directly integrated withthe conduit structure. The monolithic, integrated design allows to reducethe amount of redundant housing material, as no additional features formechanically aligning and fastening the pump housing or connecting thechannel to the pump are required.

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5.4.3 Actuator Electronics

Actuator electronics of the laboratory demonstrator consist of a DC/DCconverter patented by Noack in [106], a microcontroller, a MEMS gyroscope,and peripheral circuitry like voltage regulators (cf. [108], figure 5.1 on page 58).

Design and development of a miniaturized version of the actuator electronicsis described in this work in section 5.5 from page page 74 onwards. At leastthe components of the DC/DC converter need to sit very close to the pumphousing, in order to keep the high-current feed lines as short as possible. Thisrequires support features integrated with the monolithic conduit structure,where the pump driver PCB could be inserted and attached, which is addressedby the author in [109].

5.4.4 Manufacturing of Monolithic, Integrated Conduits

Manufacturing of monolithic, integrated conduits is difficult, due to theircomplex shape. Having only the electrode openings left for access to the insidegeometry, injection molding with lost core would be a traditional manufacturingtechnology allowing to produce this geometry. Injection molding allows touse plastics as conduit material, which fulfills the requirement of electricalisolation between the fluid, conduit walls, electrodes, and pole pieces (cf.section 5.4.2). In addition, injection molding results in good surface finish,which is important for the inside walls of the conduit in order to not increasethe flow resistance due to excessive wall roughness. However, due to themultiple molding stages required, manufacturing costs are too high to be usedin the scope of this project.

With the rise of additive manufacturing technologies over the last decades,low-cost alternatives to injection molding exist. From the diverse 3D-printingtechnologies and materials, selective laser sintering (SLS) from polyamidepowder was selected. The advantages of this combination are plenty:

– Full flexibility in conduit geometry, with wall thicknesses down to 0.7 mm.

– Low-cost procurement with short term delivery from dedicated suppliersover the Internet.

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– Similar coefficient of thermal expansion as the used metallic fluid.

– High mechanical and temperature stability of the printed parts.

While polyamide, better known as Nylon, is widely used in space applications,its hygroscopic behavior has caused catastrophic failures on space missionsdue to out-gassed and condensed water vapor, as stated by Gross in [110].The use of Nylon for space applications is therefore not recommended. WithPEEK, a more expensive, alternative material is readily available for SLSprinting. For functional verification, however, polyamide represents the betterchoice.

Additional disadvantages of using parts produced by SLS from polyamidepowder are rough surfaces and removal of residual powder from the interior ofthe channel. Solutions to those problems are addressed by the author in [109].While removal of the powder from the conduit is just a time-consuming taskthat requires manual labor, surface treatment inside the conduit is difficile.Watzinger in his Bachelor’s Thesis [111], conducted under the supervisionof the author, developed a process to improve the interior surfaces of 3D-printed polyamide parts using acids under increased temperatures. While theresults regarding surface improvements are more then promising, the longterm durability of the treated polyamide parts is severely impaired.

5.4.5 First Rapid Prototyping Experiences

A first specimen of a monolithic, integrated conduit is shown in figure 5.6.The outer dimensions of the conduit are 80 × 80 mm, and three flanges areadded for mounting onto a piece of veroboard. Conduit mass is less than 4 g,and it can hold about 18 g of liquid metal. The presented conduit is not yetequipped with mounting features for the pump electronics, as it was meantfor gaining experience with 3D-printed parts and powder removal. The mainusage is for functional demonstration on a gas bearing with the electroniccomponents mounted on veroboard (cf. sections 5.5 and 5.6).

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Figure 5.6: First monolithic, integrated fluid-dynamic actuator conduit [112]

5.4.6 Conduit Geometries for CubeSat Applications

Following functional verification of the first integrated conduit and the develop-ment of a miniaturized version of the pump driver electronics (cf. section 5.5),the conduit is adapted to be integrated with the pump driver and the satellite.Gridded support structure is added to the bottom of the actuator, which canbe seen in figure 5.7. This increases the contact area between the conduitstructure and the solar panel, to which the actuator will be glued (cf. chap-ter 6). The gridded structure at the bottom follows the design of BEESAT’sprimary structure, and therefore allows to use the same solar panel on sideswith or without pFDAs.

Additional structure was added to support the electronics board and thetransformer used with the pump driver. Thus preventing damage to thosecomponents due to mechanical loads during rocket launch. To increase satellitestiffness, the battery compartment on the BEESAT satellites is connectedwith the solar panels and the outer structure using two screws per side. Touse the same screws with the pFDAs, the conduit is further equipped with twosleeves, that later incorporate metallic spacers to have a solid mechanical andelectrical connection between the solar panels and the battery compartment.

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Figure 5.7: Monolithic, integrated fluid-dynamic actuator conduit [112]

Figure 5.8: Circular integrated conduit used for student projects

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5.4.7 Advantages of Conduit Rapid Prototyping

One of the advantages of conduit rapid prototyping is the ability to quicklydesign and produce new conduit variants. This ability was demonstrated withthe development of a circular conduit with square cross section (cf. figure 5.8).This actuator is used for education of aerospace engineering students in thescope of the spacecraft dynamics and control lecture at TU Berlin.

Design of the adapted version was done in less than one day, printed partswere shipped within one week. This shows the great flexibility of the selectedmanufacturing process and the potential for adapting conduit geometry.

5.5 Driver Electronics

Driver electronics development was carried out in three stages: First, the samecomponents used by Noack for his laboratory demonstrator (cf. [108]) wereused for functional verification of the first printed conduits on a veroboardbasis (cf. section 5.6). Second, a miniaturized version that incorporates anadditional shunt resistor in the pump power supply line was developed. Usingthis version, the pFDA was integrated with the solar panel, and environmentaltests conducted (cf. section 6.3). The same version is also used for spaceengineering education (cf. section 5.4.7). Functional verification using thissecond version showed the need to have a flexible development platform, inorder to conduct research using different supply voltages, additional sensors,different transformers, a variation of conduits, and combinations of those.

5.5.1 Electronics Miniaturization

Miniaturization of pFDA electronics was attempted straightforward, as ad-dressed by the author in [64]. Where available, smaller packages of electroniccomponents were selected. An additional shunt resistor and operational ampli-fier were introduced to the board to be able to measure pump input current(cf. figure 5.9). Solid lines depict information flow, dashed lines representpower flow to the pump. Power flow to other components is not displayed.

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5.5 Driver Electronics 75

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EMP

µC

OPA

DC/DC

SR

GYRO

IF

Figure 5.9: Block diagram of initial pFDA electronics (taken from [112])

The I2C bus on the external interface is mapped to the internal one via atunnel functionality implemented in the firmware. This allows access to allinternal functionality, including set-up and read-out of the gyroscope. This isa major improvement in comparison to earlier implementations, which allowsthe functional verification (cf. section 5.6) to be carried out using the pFDAelectronics with a lean hardware setup that comprises an additional microcon-troller to map from I2C to USART and a Bluetooth interface to communicatewith a desktop computer.

Miniaturization lead to a size of 34 × 24 mm of the pump driver PCB, whichalready includes the transformer. Overall height of the electronics board is lessthan 5 mm and height of the transformer less than 9 mm. With all componentsassembled, the pump driver weighs less than 22 g. The first fully-integratedpFDA optimized for CubeSat application is shown in figure 5.10. In additionto the microcontroller firmware running on the pFDA, functions to remotelyoperate the actuator were developed using GNU Octave [95]. This enables the

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Figure 5.10: First fully-integrated picosatellite fluid-dynamic actuator optimized forCubeSat application [109]

automated execution of predefined maneuvers on the air bearing test facilitywith parallel read out of gyroscope data and actuator telemetry. Functionalverification is build around this software library.

One drawback of the chosen implementation is, that the pump componentsuse the same supply voltage as the logic part of the electronics. This disablesthe possibility to provide higher voltages to the pump in order to increase theoutput torque and angular momentum storage capacity of the actuator.

5.5.2 Flexible Development Platform

During functional verification of the fully miniaturized pFDA, the need to usedifferent input voltages and therefore different electronic components arose.The situation was aggravated, when the gyroscope used for the miniaturized

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5.5 Driver Electronics 77

Figure 5.11: Components of the flexible development platform

Figure 5.12: Flexible development platform assembled for air bearing use

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pump driver was discontinued during development of the circular actuator (cf.section 5.4.7). Discontinuation of the gyroscope lead to a redesign of thepump driver PCB, and major changes in the software. In support of futuredevelopment of pFDA conduits, electronics, and software, a flexible platformbased on breadboards was created.

Figure 5.11 shows a selection of components of this development platform,like DC/DC converters (left), conduits (center), breadboard with electronics(upper right), and a mass dummy required for angular momentum estimation(lower right). Breadboard and conduit boards are stackable (cf. figure 5.12),and conduit electrodes are connected to the converters using screw terminals.This allows for a flexible combination of components for gas bearing tests.Moreover, the breadboard enables the use of high-side current monitor break-out boards to measure pump and logic component power demand duringactuator operation on the gas bearing. Hence, power demand monitoring iseasily synchronized with angular rate measurements.

5.6 Functional Verification

Like for the solar antenna (cf. chapter 3), no environmental verification is con-ducted on component level. Environmental tests are conducted after assemblywith the multi-functional solar panel (cf. section 6.3). PFDA functionalverification was carried out in multiple stages in parallel to conduit develop-ment and pump driver miniaturization. Following design and procurementof the first 3D-printed conduits (cf. section 5.4.5), the demonstrator shownin figure 5.13 was assembled and dynamic properties measured on the airbearing. Measurement results were first published by the author in [112], andshowed their potential for creating large amounts of torque. However, it wasnot possible to achieve the same levels of torque and angular momentum, thatNoack presented in [108]. Measurements further show a perturbation torquecreated by the air bearing that influences the measurement results.

For functional verification of the first fully integrated actuator (cf. sec-tion 5.5.1), measurement frequency is increased from 2 Hz to 10 Hz in orderto resolve time behavior of the acceleration phase. An improved test sequence

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5.6 Functional Verification 79

Figure 5.13: First demonstrator of a pFDA using a 3D-printed conduit [109]ii

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0 5 10 15 20 25 30 35 40 45 50-1.2

-0.9

-0.6

-0.3

0

0.3

0.6

0.9

1.2

time 𝑡 [s]

angu

larr

ate

𝜔𝑧

[°/s]

20 %40 %60 %80 %100 %

Figure 5.14: Angular rate measurements of a pFDA, adapted from [64]

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is applied: pump speed is commanded in discrete steps between −100 % and100 % with intermittent phases of no actuation. This sequence allows to filtergas bearing disturbance from angular rate measurements. Figure 5.14 showsfiltered angular rate measurements for steps of 20 % in both directions.

The actuator has symmetric angular rate for positive and negative directions,and shows linear increase of angular rate with increasing pump speeds. Theminiaturized driver electronics reach the same order of magnitude of angularrate as the initial 3D printed demonstrator. This is however not enough tobe used on a 1 U CubeSat mission for three-axis attitude control. There, theinertia of the full satellite is much larger than that of the air bearing, andtherefore satellite angular rates will be reduced.

5.6.1 Power Consumption

Figure 5.15 shows the pump power demand at a supply voltage of 5 V overpump speed and pump frequency. Minimum power required to let the pumprun at zero speed is less than 10 mW. For frequencies below 10 kHz, maximumpower is less than 40 mW, and power increase develops roughly linear withcommanded pump speed. At higher frequencies, maximum power demand isgreater than 40 mW, and power increase is better approximated by a secondorder polynomial. For actuator operations on a spacecraft, the respectivepower consumption development over commanded speed needs to be takeninto account.

Similarly, electric power consumption for 12 V supply voltage is shown infigure 5.16. Minimum power consumption for idle pump is 26 mW in this case.The contour plot shows a similar behavior as for 5 V, where for frequenciesbelow of 10 kHz the maximum power is reduced. For frequencies of 11 kHzand above, less than 190 mW of electric power are required by the pump.

5.6.2 Dynamical Properties

Due to data rate limitations and power supply considerations, actuator dy-namical properties are verified using the flexible development platform (cf.

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-100 -75 -50 -25 0 25 50 75 1006

8

10

12

14

16

speed [%]

frequ

ency

[kH

z]

10 20 30 40power [mW]

Figure 5.15: Pump power demand at 5 V in mW

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-100 -75 -50 -25 0 25 50 75 1006

8

10

12

14

16

speed [%]

frequ

ency

[kH

z]

50 75 100 125 150 175power [mW]

Figure 5.16: Pump power demand at 12 V in mW

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section 5.5.2). Angular rate measurements displayed in figure 5.14 show, thatthe step response of the actuator on the air bearing has first-order, linear time-invariant (LTI) system characteristics. To better estimate torque and angularmomentum of the pFDAs, a model-based parameter estimation approach isimplemented.

5.6.2.1 Model-Based Parameter Estimation

The step response of the actuator is that of a first-order, LTI system, which isdefined for the angular rate as

𝜔 (𝑡) = 𝜔max(︁

1 − 𝑒− 𝑡𝑇

)︁, (5.30)

where 𝜔 (𝑡) is the angular rate over time 𝑡, 𝜔max is the angular rate atsaturation, and 𝑇 is the time constant. Maximum angular acceleration

�̇�max = 𝐾

𝑇. (5.31)

Figure 5.17 shows a measurement of angular rates �̂� that is cleared from long-period disturbances. Estimated parameters 𝜔max and 𝑇 are shown, togetherwith the derived angular rate 𝜔 (𝑡). Steady state behavior displays somefluctuations around 𝜔max. Those are created by small roll and pitch motionsof the test platform on the air bearing due to gravitational disturbances.

The advantage of using a model-based parameter estimation approach is, thatthe estimation of angular acceleration �̇�, which is proportional to torque 𝜏 (cf.section 5.6.2.3), becomes more reliable. Estimating the angular accelerationbased on a linear least-squares fit of the first 𝑛 values on the slope shown infigure 5.17 would lead to an underestimation of the initial torque. To furtherimprove accuracy of the estimated parameters, parameter estimation needs tobe applied to a batch of multiple, similar measurements.

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0.25

0.5

0.75

1

1.25

1.5

𝑇

𝜔max

time [s]

angu

larr

ate

[°/s]

�̂�

𝜔 (𝑡)𝜔max�̇�max

Figure 5.17: Actuator parameter estimation based on a first-order LTI model

5.6.2.2 Automated Parameter Estimation

For a batch of eleven measurements, estimated and derived parameters areshown in figure 5.18. For all parameters, calculated mean, median, as well asminimum and maximum are displayed.

Estimated angular rate shows little deviation of minimum and maximumvalues (cf. figure 5.18(a)). Mean and median values have good correlation.The number of data points is sufficient for a good gain estimation. Timeconstant estimation shows large deviation of minimum and maximum values(cf. figure 5.18(b)). The median starts to deviate from the mean for increasingpump speeds. This is a hint, that not enough data points are available fortime constant estimation, and additional error is introduced for increasingpump speeds.

Data presented in figure 5.18 is used to calculate maximum angular accelerationaccording to eq. (5.31) (cf. figure 5.18(c)). The plot shows a linear dependencyof maximum angular acceleration on commanded pump speed.

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25 50 75 1000

1

2

3

4

5

6

7

pump speed [%]

𝜔max [°/s]

meanmedianminmax

(a) angular rate

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25 50 75 1000.4

0.45

0.5

0.55

0.6

0.65

0.7

0.75

pump speed [%]

𝑇 [s]

meanmedianminmax

(b) time constant

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25 50 75 1002

4

6

8

10

12

14

16

pump speed [%]

�̇�max [°/s2]

meanmedianminmax

(c) angular acceleration

Figure 5.18: Estimated actuator parameters

In order to calculate actuator angular momentum 𝐻 and torque 𝜏 fromangular rate 𝜔 and �̇�, respectively, knowledge of mass moment of inertia 𝐼

of the complete hardware setup floating on the gas bearing is required (cf.eqs. (5.2) and (5.24)). Especially for the flexible development platform (cf.section 5.5.2), estimating the inertia from e.g. computer aided drawing (CAD)software is impossible.

5.6.2.3 Actuator Dynamic Properties

Under the assumption, that the actuator always creates the same angularmomentum for a commanded pump speed, two systems incorporating thisactuator are linked by

𝐻1 = 𝐻2, (5.32)

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5.7 Fluid-Dynamic Actuators and Reaction Wheels 85

where 𝐻1 and 𝐻2 is the angular momentum of the first and the secondsystem, respectively. This is equivalent to

𝜔1 · 𝐼1 = 𝜔2 · 𝐼2, (5.33)

where 𝜔1, 𝜔2 and 𝐼1, 𝐼2 are measured angular rates and inertias of systemsone and two, respectively. If the inertia of the second system is the same asthat of the first one except for a known difference 𝐼z, this may be written as

𝐼2 = 𝐼1 + 𝐼z. (5.34)

Inserting eq. (5.34) into eq. (5.33) and solving for 𝐼1 leads to

𝐼1 = 𝜔2𝜔1 − 𝜔2

· 𝐼z. (5.35)

In order to estimate the gas bearing inertia, two batches of measurements,where the second batch is conducted using the additional inertia 𝐼z, arethen processed according to section 5.6.2.2. The measured angular rates𝜔1 and 𝜔2 are then used together with 𝐼z to find an inertia of the firstsystem of 𝐼1 = 907 kg mm2. This leads to a maximum angular momentum of𝐻max = 83.8 µN m s and a maximum torque of 𝜏max = 171 µN m.

5.7 Fluid-Dynamic Actuators and Reaction Wheels

The smallest, currently available reaction wheel for CubeSats is the AstrofeinRW 1, which comes in two different types (cf. [113]). While identical in sizeand power consumption, the two types have different flywheel inertias, andtherefore different angular momentum and torque. To compare those tworeaction wheels against the pFDAs, they are considered to be used on a singleunit CubeSat with 1.33 kg maximum mass, and an edge length of 10 cm (cf.[25]). Assuming homogeneous mass distribution, satellite inertia about one ofits axes

𝐼1U = 1.33 kg · 100 cm2

6 = 2.22 g m2. (5.36)

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Using eqs. (5.2) and (5.24) with above inertia and dynamical properties ofRW 1 and pFDAs, angular rate and angular acceleration are calculated andshown in table 5.1 for all three of them. Data in that table points out, thatpFDAs are capable of producing slightly less maximum angular rate as the RW1 type B, about 2.17 °/s. In comparison to single unit integrated platforms,this is found to be in a good range of maximum angular rate for single unitCubeSats (cf. table 2.2 on page 16).

Table 5.1: Comparison of dynamical properties of CubeSat reaction wheels and thepicosatellite fluid-dynamic actuator

Actuator 𝜔 �̇�

°/s °/s2

RW 1 A 15 0.595RW 1 B 2.58 0.103pFDA 2.17 4.42

5.7.1 Angular Rate and Acceleration

To better understand the implications of using pFDAs for satellite attitudecontrol, simulated data for angular rate 𝜔 and angle 𝜙 for RW 1 types Aand B as well as the pFDA is shown in figure 5.19. Reaction wheel angularmomentum saturates after 25 s, while the pFDA, owing to its first-order stepresponse, needs only about 1.5 s to reach 95 % of its maximum angular rate.While RW 1 type B travels 32.3° until saturation, the pFDA travels 53° inthe same time. Even up to an angle of 13.3°, reached after 6.70 s, the pFDAtravels faster than RW 1 type A. If all actuators are kept at the saturatedlevels, the reaction wheels would travel further with constant angular rate,and linearly growing angles.

5.7.2 Time-Optimal Slew Maneuvers

A more typical scenario for reaction wheel applications are time-optimal slewmaneuvers, where the wheels are accelerated up to the point of angular

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5.7 Fluid-Dynamic Actuators and Reaction Wheels 87

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2.5

5

7.5

10

12.5

15

time [s]

angu

larr

ate

[°/s]

𝜔rw1,A𝜔rw1,B𝜔pFDA

(a) Angular rate time curve

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30

60

90

120

150

180

time [s]an

gle

[°]

𝜙rw1,A𝜙rw1,B𝜙pFDA

(b) Angle time curve

Figure 5.19: Comparison of simulated actuator properties

momentum saturation, and then decelerated up to the point of zero angularrate. Simulated results for time-optimal maneuvers using RW 1 types A andB and a saturated maneuver using the pFDA is shown in figure 5.20. Due tothe short acceleration and deceleration phases of the fluid actuator, it travels107° in 50 s, while the RW 1 type B travels only 64.6°. RW 1 type A with372° traveled in 50 s is completely out of range for the other two actuators.Please note, that the extremely short deceleration phase of the pFDA resultsfrom commanding the maximum angular rate in the opposite direction, andswitching the actuator off close to zero-crossing. This results in a decelerationtime span of less than 0.4 s.

5.7.3 Analysis of Traveled Angles

In order to finally compare reaction wheel to pFDA agility, simulations of bothtime-optimal and time-non-optimal maneuvers are run, and the traveled angleis plotted over the total maneuver time (cf. figure 5.21). In this sense, amaneuver is considered time-optimal, if the total maneuver time is less thanor equal to twice the angular momentum saturation time of 25 s. Only then

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0 10 20 30 40 500

2.5

5

7.5

10

12.5

15

17.5

20

time [s]

angu

larr

ate

[°/s]

𝜔rw1,A𝜔rw1,B𝜔pFDA

(a) Angular rate time curve

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0 10 20 30 40 500

20

40

60

80

100

120

140

160

time [s]

angl

e[°]

𝜑rw1,A𝜑rw1,B𝜑pFDA

(b) Angle time curve

Figure 5.20: Comparison of simulated time-optimal slew maneuvers

does the maneuver consist of an acceleration phase directly followed by adeceleration phase. Time-non-optimal maneuvers are then all maneuvers, thatincorporate a phase of constant angular momentum between the accelerationand deceleration phases. The border between the two regimes is marked infigure 5.21 with an extra tick on the top x axis. Over the observed range ofup to 160 s, the pFDA exposes linear behavior with a slope of 2.17 °/s. RW 1type A shows only parabolic behavior, as the y axis is cut-off at 360°. AndRW 1 type B shows parabolic behavior until angular momentum saturation,and from then on linear behavior with a slope of 2.58 °/s (cf. table 5.1).

Simulation of traveled angles shows two intersections between pFDA andreaction wheel trajectories. PFDA and RW 1 type A intersect at 13.6 s anda traveled angle of 27.9°. Hence, the pFDA is more agile than this reactionwheel for traveled angles smaller than this limit. Intersection between pFDAand RW 1 type B happens far in the non-optimal region at 151 s and an angle

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0 20 40 60 80 100 120 140 1600

60

120

180

240

300

36013.6 50 150

27.9

324

time [s]

angl

e[°]

Δ𝜑rw1,AΔ𝜑rw1,BΔ𝜑pFDA

Figure 5.21: Maximum angles traveled using time-optimal and non-optimal maneu-vers

of 324°. Here, the pFDA shows that it is far superior to the low angularmomentum version of the reaction wheel.

5.8 Redundancy Concepts

Due to their large dimensions in the 𝑥𝑦 plane, planar actuators are out of scopefor redundancy concepts featuring four actuators with linearly dependent axesof angular momentum. From the conduit geometries derived in section 5.3,this leaves the three-dimensional conduits from section 5.3.2 as only options.

5.8.1 L-Shaped Conduits

Using four L-shaped conduits allows to arrange them in a way, that a tetrahe-dron configuration of angular momentum vectors is achieved (cf. figure 5.4).

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Figure 5.22 shows this configuration together with the direction of angularmomentum vectors 𝐻1 through 𝐻4 of all four actuators. The four angularmomentum vectors span the tetrahedron, eponymous for the configuration.Conduit width 𝑐 is set to the available distance inside the CubeSat structure.Setting 𝑎 = 𝑐

2 and applying eqs. (5.17) and (5.18) for an angle

𝛽 = tan−1(︂

1√2

)︂(5.37)

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𝐻1

𝐻2

𝐻3

𝐻4

Figure 5.22: L-shaped conduits in a redundant configuration

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leads to the tetrahedron configuration with angular momentum

𝐻 = 𝑣𝜌𝑆𝑐2√3. (5.38)

During nominal operation of all four actuators, this would then lead tomaximum angular momentum

𝐻max = 2𝑣𝜌𝑆𝑐2√3. (5.39)

The advantage of this arrangement is, that the conduits now fit in verywell with cuboid envelopes of CubeSat subsystems. Additional mechanicalfeatures allow to replace tertiary satellite structure with enhanced, L-shapedconduits. Sharing mass budget between ADCS and structures increasesthe mass available for higher performance payloads. Figure 5.23 shows afirst experimental assembly of a redundant pFDA configuration based onL-shaped conduits. Using an adapter structure for the flexible developmentplatform on the air bearing (cf. section 5.5.2), it is possible to conductfunctional verification of a single L-shaped pFDA. The adapter aligns the

Figure 5.23: Experimental assembly of redundant fluid-dynamic actuator configura-tion based on L-shaped conduits

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angular momentum vector of the actuator with the axis of rotation of theair bearing. Conduction of functional verification of the fully assembled,redundant configuration on the air bearing is impossible due to residualdisturbance torques caused by improper balancing of the center of mass.

In the scope of the TUPEX-6 project, student team members develop apseudo-CubeSat, that will be deployed from a REXUS sounding rocket in April2019 [114]. The REXUS/BEXUS program is realized under a bilateral AgencyAgreement between the DLR and the Swedish National Space Board (SNSB).The Swedish share of the payload has been made available to students fromother European countries through the collaboration with the European SpaceAgency (ESA). The author of this work contributes design, development, andfunctional verification of the L-shaped pFDAs. The pseudo-CubeSat, dubbedfree-falling unit (FFU), including all relevant subsystems, and the deployeron the sounding rocket are contributed by the student team. Experimentobjective is the functional verification of redundancy strategies in a reducedgravity environment.

Among secondary mission objectives is demonstration of increased payloadcapacity of the redundant set of actuators being considered. After deploymentfrom the sounding rocket and conduction of the core experiment, the FFUwill fall from an altitude between 80–90 km. A parachute will be used fordecelerating the downward velocity and to allow for soft landing. Stored insidethe FFU, the parachute will occupy a volume of about one third of the overallsatellite volume. Additionally, one of the PCBs shown in figure 5.23 will hostthe so-called recovery system electronics, which include GPS based locationand RF broadcasting services.

5.8.2 Crown-Shaped Conduits

The geometry of the crown-shaped conduit (cf. figure 5.5) allows to use fouractuators in a rotation-symmetric configuration at one end of a CubeSat (cf.figure 5.24). A tetrahedron configuration is achieved by setting the anglebetween the reference plane and the angular momentum vectors to the samevalue as for the L-shaped conduit. According to eq. (5.23), this is achieved

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by setting 𝑎 = 𝑏 = 𝑐. For this choice, magnitude of the angular momentumvector is given as

𝐻 = 𝑣𝜌𝑆𝑐2√3, (5.40)

with maximum angular momentum during nominal operation of all fouractuators given by

𝐻max = 2𝑣𝜌𝑆𝑐2√3. (5.41)

Comparison of eqs. (5.38) and (5.40) shows, that for a single unit CubeSat,the crown-shaped actuator setup is only capable of producing a quarter ofthe torque of the l-shaped actuator setup, if the same fluid velocity, density,and cross-sectional area are assumed for both. In order to increase angularmomentum, a more powerful pump, a different fluid, or a larger cross-sectionalarea are required.

The advantage of the crown-shaped setup is that the actuators are formingan opening to one side, which could accommodate subsystem PCBs or anoptical payload with a larger focal length.

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𝐻1

𝐻2

𝐻3𝐻4

Figure 5.24: Crown-shaped conduits in a redundant configuration

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6 A Highly Integrated Single Unit CubeSatSolar Panel

The CubeSats with the best payload volume ratio, Planet Labs’ triple unitDove satellites, integrate the bus components around parts of the instrumentin the rear sixth of the spacecraft, leaving a large part of the satellite availablefor the integration of a high performance optical payload (cf. section 2.1.1).Payload-to-satellite volume ratio of this system is unprecedented, and none ofthe existing integrated platform solutions for single to triple unit CubeSatsmatches this ratio (cf. section 2.1.2). Integrated platforms for single unitCubeSats, which were found to be the most relevant satellites for universityscience and Earth observation missions (cf. section 1.1.3), achieve only up to25 % of payload volume ratio.

Bad volumetric efficiency of single unit platforms is due to multiple factors:

– CubeSat vendors base their platform designs on existing subsystemswhich are not optimized for volumetric utilization, but aiming for multi-purpose, multi-customer use cases (cf. pages 12 to 16, [30]).

– Based on the PC/104 form factor, those subsystems are inept to realizea high integration density, as the standard accounts for single-purpose,multi-vendor satellite systems (cf. section 2.1.2.7).

– Due to the use of a multitude of different connectors and cables be-tween the subsystems, additional volume is occupied by harness (cf.section 2.1.2.7).

In order to increase integration density of single unit CubeSats to supporthigh performance payloads, subsystems need not only be optimized in termsof their characteristics. During development of the TUBiX20 nanosatelliteplatform, major design requirements were reusability, adaptivity, and robustness.Modularity and extensive reuse of hard- and software on various levels were

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96 6 A Highly Integrated Single Unit CubeSat Solar Panel

found to be key to meet those requirements, as stated by Barschke and Gordonin [115]. Applying a similar development approach to CubeSat design allowsto develop a single unit platform where the entire system is considered foroptimization and offers more resources to higher performance payloads.

On the hardware layer, TUBiX20 achieves modularity using common mechani-cal parts in the primary structure, an electronics box with a PCB backplane,and subsystem boards that all feature the same interface for data and electricpower [115]. A similar design based on a backplane and a common inter-face is used for the UWE-3 satellite and upcoming missions developed atJulius-Maximilians-Universität Würzburg [116]. While this concept has flightheritage on small satellite missions, it is inefficient in terms of volumetricutilization (cf. section 2.1.3.2).

Total miniaturization and integration of all satellite subsystems on one singlePCB, currently investigated in the scope of the the PiNaSys II project atTU Berlin (cf. section 2.1.3.3), is opposite to modular spacecraft designapproaches. PiNaSys II aims at demonstrating a maximum integration densityon a quarter unit CubeSat bus, which is therefore only capable of hostinglower performance payloads.i

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Figure 6.1: Schematic view of a proposed high density single unit CubeSat usinghighly integrated, multi-functional solar panels

Obviously, neither strictly modular nor strictly integrated approaches to singleunit CubeSat design are able to maximize payload performance over multiple

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97

missions in an university environment. A concept, where subsystem function-ality is distributed between a small set of densely integrated PCBs in whatresembles a classic CubeSat board stack, and up to six highly integrated,multi-functional solar panels was presented by the author in [63]. Using thissetup, the satellite bus is wrapped around a larger and better compactedpayload volume at the spacecraft core. Figure 6.1 gives a schematic impressionof such a proposed high-density, single unit CubeSat.

Spacecraft subsystems are condensed to an OBC board (blue), a PCDU board(red), a battery compartment (gray), and up to six solar panels (green). Thelarge, compact payload volume (yellow) is shown in the center. For better visi-bility, three side panels as well as the primary structure are hidden. Consideredgeometrical dimensions of the subsystems take into account recent develop-ment for current and upcoming CubeSat missions at TU Berlin. Therefore,payload volume shown in figure 6.1 promises to achieve a payload volumeratio of up to 60 %. As the schematic view neglects the volume required forharness and pFDAs, the actual number might be lower than this.

Based on the TUBiX20 node concept (cf. [115]), a conceptual representationof a proposed bus architecture is given in figure 6.2. The nodes displayeddo not follow the same strict distinction between logical nodes (functionalunits) and physical nodes (specific microcontrollers of a node) as defined byi

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PCDU OBC solarpanel

payload

power

data

Figure 6.2: Nodes of a proposed high density single unit CubeSat

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Barschke and Gordon in [117]. Thus, for example, each PCDU or OBC logicalnode could, but not necessarily needs to, consist of multiple physical nodes.

PCDU, OBC, and payload nodes are located in the board stack, while thesolar panel nodes are mounted to the CubeSat structure. PCDU operates in ahot redundant worker/monitor configuration, and is the redundancy controlauthority for cold redundant OBC and solar panel nodes. Like the TUBiX20platform and the BEESAT-5 to BEESAT-8 spacecraft, the data bus is basedon a redundant CAN interface. The power bus uses dedicated latched andmonitored power lines to distribute electric power to the nodes.

Design of PCDU, OBC, and payload nodes is not in the scope of this thesis.Backed by the integration density advancement of BEESAT-5 to BEESAT-8,all components required for a fully functional CubeSat bus not located onthe solar panels are assumed to be found on either the PCDU or OBC. Thelast chapter of this work focuses on integration of the solar antenna (cf.chapter 3), an optimized magnetic actuator (cf. chapter 4), a square pFDA(cf. chapter 5), and additional subsystem functionality into a multi-functionalCubeSat solar panel.

6.1 Current Integration Density of CubeSat Solar Panels

The primary function of CubeSat solar panels is to host solar cells on theoutside of the satellite structure and connect them to the power control anddistribution unit, which is commonly found in the board stack that makes upthe main satellite bus. A growing number of panels is equipped with additionalcomponents and functionality, like sun sensors, magnetic coils, or maximumpower point trackers. To compare characteristics of body-mount single unitCubeSat solar panels currently in use, their properties and components aregathered in tables 6.1 and 6.2. The upper part of each table holds panelsoffered by CubeSat component vendors, the lower part lists a selection ofcustom made panels which are relevant in the scope of this work. The lists donot claim for completeness, as many university and also commercial CubeSatsfeature custom made panels.

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6.1 Current Integration Density of CubeSat Solar Panels 99

6.1.1 Mechanical Properties

CubeSat solar panels are based on PCB, aluminum sheet, or carbon-fibercarriers onto which solar cells and electronic components are mounted. Ad-vantages of using PCBs as carrier are, that multiple copper layers and vias canbe used to conduct the solar cell current, all electronic components can besoldered directly in place, and boards can be procured at low cost. Aluminumsheet and carbon fiber carriers require additional thin-film conductive layers,increasing complexity and cost of the panels. Aluminum and carbon-fibercarriers, however, can carry higher structural loads than PCBs. Therefore,utilization of such boards can reduce the mass of CubeSat primary structure.

All single unit solar panels have maximum dimensions of 100 × 83 mm, drivenby the requirement to fit between the CubeSat rails (cf. table 6.1, [25]). Thede-facto standard for PCB carrier thickness is 1.6 mm. If magnetorqers areembedded into the carrier board, the thickness might deviate from the standard.For aluminum or carbon-fiber based panels, thicknesses are commonly kept atthe same value in order to stay compatible with the standardized frames andCubeSat kits of major vendors. Masses of the panels are between 30–60 g (cf.table 6.1). CubeSat designs with six panels therefore need to reserve about20–25 % of total satellite mass for solar panels, if heavier panels are required.

Table 6.1: Single unit CubeSat solar panel properties

Supplier Name Mechanical Solar Cells

𝑀 𝑙 𝑤 ℎ No Eff. 𝑃

g mm mm mm # % W

Clyde Space 25-02869 n/a n/a n/a 1.6 2 28.3 2.1 [118]DHV Tech. SPC-CS10 39 n/a n/a n/a 2 30 2.4 [119]EnduroSat XY /MTQ 50 98 82.6 2 2 29.5 2.4 [81]GomSpace P110UA 57 98 82.6 1.6 2 30 2.3 [80]ISIS 50 98 82.6 1.6 2 30 2.3 [120]nano avionics 1U-S 35 100 82 1.6 2 28.7 2 [121]nano avionics 1U-SMS 40 100 82 1.6 1 28.7 1 [121]

Poli. Torino CubePMT n/a 98 82.5 1.6 2 26 n/a [93]Uni Würzburg UWE-3 [122]TU Berlin BEESAT 34.4 98 82 1.6 2 29.6 2.4 [54]

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Most solar panels feature mounting holes in the four corners of the panel.Presence of the holes along the panel’s outer perimeter reduces the enclosedarea for embedded air coils and hence the maximum achievable magneticdipole. To avoid this loss, BEESAT solar panels feature two mounting holesclose to their center.

6.1.2 Power Generation

Usually, two large-area, triple-junction solar cells with 28 % and above efficiencyare used for generating about 2 W power per single unit solar panel (cf.table 6.1). Thus, panel surface area is covered more or less completely bysolar cells. Components like precision sun sensors, patch antennas, RBF pins,umbilical connectors, and others reduce the available surface area. This eitherleads to a reduction in number of large-area solar cells, or usage of a largernumber of small-area cells. Due to their electrical properties, combinations ofdifferent shapes or types of solar cells on a single panel are usually avoided.

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(a) Same cell orientation with mounting holesin the corners

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(b) Opposite cell orientation with mountingholes in the center

Figure 6.3: Solar cell orientation and mounting hole placement

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For mounting large-area solar cells, two general methods are applicable. Ifcells are oriented in the same direction with respect to the chamfered side(cf. figure 6.3(a)), electrical connection between cells is simplified, but freesurface area is reduced to a narrow brim at the panel’s edges and a narrowgap between the cells. Hence, additional components on the outside need tohave a small footprint area.

If chamfered sides of the cells are pointing in opposite directions, electricrouting is less simple, but a significant gain in connected area for additionalcomponents is observed (cf. figure 6.3(b)). Precision sun sensors or RBF pinsmight be located between the solar cells while mounting holes are sitting atthe chamfered cell edges. Yet, inter-cell area is not large enough to host patchantennas for high-throughput communication (cf. chapter 3).

Apart from solar cells, the two bottommost entries in table 6.2 feature MPPTcapabilities implemented in hardware and software on the panel. Panels notfeaturing built-in MPPT capabilities have to rely on this being provided by thePCDU. Busch in [56] states that having MPPT implemented directly on thepanel, "optimal performance can be achieved for each panel separately" whilesingle panels are exposed to different irradiation and temperature conditions.

6.1.3 Attitude Determination Sensors

Simple Solar panels feature photodiodes acting as coarse sun sensors, whichrequire temperature sensors to compensate thermal offsets (cf. table 6.2).The advantage of using photodiodes is their small size. Precision sun sensors,e.g. based on position-sensitive devices (PSDs) or CMOS technology, requiremore electrical power and more area on both sides of the panel.

MEMS magnetic field sensors and gyroscopes, usually located on ADCS boards,are rarely found on commercial CubeSat solar panels (cf. table 6.2). Thosedevices exhibit a pronounced temperature dependency of magnetic field andgyroscopic measurements. Their application on solar panels, which experiencethe largest temperature swings during satellite life time, requires additionaleffort for temperature calibration. Placing magnetic field sensors close tocomponents carrying large currents, like solar cells or magnetorquers, further

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Table 6.2: Single unit CubeSat solar panel components

Supplier Name Sensors Actuators

Sun Temp. Magn. Gyro TorquermA m2

Clyde Space 25-02869 X X – – X 80 [118]DHV Tech. SPC-CS10 X X X – – – [119]EnduroSat XY /MTQ X X – X X 131 [81]GomSpace P110UA X X – – X 38 [80]ISIS X X – – – – [120]nano avionics 1U-S X X – – X 70 [121]nano avionics 1U-SMS X X – – X 70 [121]

Poli. Torino CubePMT X X X X X n/a [93]Uni Würzburg UWE-3 X X X X X 28 [122]TU Berlin BEESAT X X – – X 46 [54]

impairs sensor readings. Using a microcontroller local to the solar panel forcorrection of temperature effects and filtering of sensor data is advantageousfor precise attitude determination [56].

6.1.4 Attitude Control Actuators

Owing to their simplicity, magnetorquers are considered to be the best choicefor coarse CubeSat pointing. Most vendors offer their panels with the optionfor an embedded air coil (cf. table 6.2 and section 4.1). Some vendors offeroptional connectors for additional, customized magnetorquers on the panel.

The advantage of using embedded air coils is that they do not occupy additionalvolume inside the satellite and there is no reduction in available area for placingelectronic components on the solar panel back. Disadvantages are the reducedarea for placing vias, high manufacturing costs for multi-layer boards, and thewidely-held assumption of small provided magnetic dipoles (cf. section 4.1).

Advantages of board-external magnetorquers are the unrestricted choice ofsolar panel carrier material, reduced production cost for two or four-layerPCBs, and flexibility in terms of late magnetic actuator selection. Also,magnetorquers and solar panels could be sourced from different suppliers. Thebiggest among the disadvantages of using board-external magnetorquers is the

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increased volumetric consumption of this solution. This effect is less severeusing wound air coils, as they can be attached to the primary structure or thesolar panels. If they are attached to solar panels, available area for electroniccomponents on the rear of the panel is though reduced (cf. section 4.1).

Except for folding mechanisms of deployable solar panels or antennas, otheractuators located on the single unit CubeSat panels are unheard of. In thescope of the AraMiS project, attaching miniaturized reaction wheels to thesolar panels, so-called intelligent tiles, of larger satellite was proposed bySperetta in [123].

6.1.5 Harness

CubeSat component vendors want their solar panels to be compatible withnot only their products, but also with those of other suppliers. This leadsto a situation, where every functionality present on a panel features its ownconnector and harness. Solar cells are connected to the PCDU, attitude sensorsand actuators to the ADCS, and so on. Therefore, volumetric utilization rateinside a CubeSat is further reduced if the number of different functionalitieswere to be increased on traditional CubeSat solar panels, as each would requireits own harness and connector.

Custom solar panels may differ from the harness scheme described above.In case of the BEESAT satellites, IDC connectors and flat ribbon cablesare used to connect all functionality of the solar panels to the satellite bus.This increases the volumetric utilization and eases assembly of the satellite.However, using flat ribbon cables is challenging, when different signals andsupply voltages have different targets to be routed to, as a connector can onlybe connected to one subsystem board. Splitting one end of the cable andusing more than one connector on that end allows to connect a single panelto multiple target systems at slightly increased cost and assembly effort.

Modern consumer electronic products heavily rely on flexible flat cable (FFC)to connect components. Multiple metallic conductors are bonded to a flexibleplastics film to form an ultra thin cable. FFCs offers great flexibility in CubeSatharness design, as the harness might be specifically adapted to the satellite

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structure and feature several connectors, both mid-cable and at the multipleends, as stated by Karuza in [124]. Disadvantages are high production costand long lead times.

One exception in terms of harness are the UWE-3 solar panels, which usedirect board-to-board connectors between panels and the backplane (cf. [56]).This is the most simple and cost effective way of connecting the panels. Theuse of a backplane, however, is inefficient in terms of volumetric utilization(cf. section 2.1.3.2).

For a modular university CubeSat platform, using FFC as harness is notapplicable. Individual connectors and cables are also not recommended forhigh performance spacecraft due to the poor volumetric utilization they impose(cf. [6]). Flat ribbon cable and IDC connectors are considered to be the bestchoice. They allow to have multiple connectors mid-cable, and if a commonbus is used to connect all subsystems (cf. section 2.2.4), they enable flexibleand modular CubeSat design with very good volumetric utilization.

6.1.6 Conclusion

Discussing the integration density of single unit CubeSat solar panels shows,that different qualities of panels exist. On the lowest level, solar panels havesimplified designs in order to achieve compatibility between products of differentvendors and to be low cost. Basic equipment of those solar panels are solarcells, coarse sun sensors, temperature sensors, and optionally magnetorquers.Harness is realized using multiple connectors, one per embedded functionality.

On the intermediate level, panels are equipped with additional attitude determi-nation sensors. BEESAT solar panels, for example, are considered intermediatelevel panels. They feature a precision sun sensor based on a PSD, with eval-uation of sensor raw data and execution of the attitude control algorithmsbeing performed by the ADCS software located on the OBC board.

High level panels incorporate precision attitude sensors, magnetorquers, PCDUfunctionality, and a microcontroller to execute those tasks locally on thepanel. One high level example is the UWE-3 panel that features attitudesensor filtering and MPPT on a local microcontroller (cf. [56]). A second

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6.2 Design of a Highly Integrated, Multi-Functional Solar Panel 105

example is the so-called "intelligent tile" dubbed CubePMT developed in thescope of the AraMiS project at Politecnico Torino [93]. This panel has noprecision sun sensor, but includes a gyroscope, a magnetic field sensor, amagnetorquer, MPPT, and voltage regulators for local power supply to theelectronics components. It is the panel with the highest integration densityup to date, but still awaits IOD. Due to the use of six different connectorsand some taller components like gyroscopes and buck-boost converters, thepositive effect on overall volumetric utilization for the satellite are believed tobe less than optimal. Notable about the CubePMT is the envisaged layout ofthe proposed AraMiS-C1 single unit CubeSat, where the satellite bus is madefrom four CubePMT tiles and two communication tiles, which leaves most ofthe internal volume available for payloads.

6.2 Design of a Highly Integrated, Multi-Functional SolarPanel

Based on preliminary work in chapters 3 to 5 it is the primary objective inthis chapter to design a highly integrated, multi-functional solar panel, takingthe considerations of section 6.1 into account. A CubeSat mission wouldthen require a single solar antenna, three pFDAs, and a set of six solar panelsin total. As a second objective, the layout of the solar panel PCB shouldtherefore allow for different configurations:

Basic: Does not include the pFDA or the solar antenna. The full area on thefront of the panel is available for two large solar cells.

Advanced control: Integrate only the pFDA with the basic version of thepanel. This requires electrical and mechanical interfaces between actua-tor and panel.

Communication and advanced control: Integrate both the solar antennaand the pFDA on the panel. Integration of the solar antenna requiresthe use of large and small solar cells.

In order to allow the three configurations to use the same PCB layout,additional care has to be taken in designing the contact surfaces for the

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different solar cells and antenna on the front of the panel. Less effort isrequired on the back, if the mechanical interface between actuator and panelis kept similar to the existing interface between BEESAT panels and structure.The electrical interface needs to be implemented in a way that allows forrobust connection.

6.2.1 Communication

The only component related to the communication system is the solar antenna(cf. chapter 3). It is located in one corner of the solar panel, and glued tomatching contact areas using conductive adhesive (cf. page 31). DC currentfrom the solar antenna is routed to the panel using two electrical contacts onthe edges of the antenna (cf. figure 3.7). A straight miniature SMA connectoris used to feed the RF signal on the back of the solar antenna. To install theantenna on the solar panel, a small opening needs to be provided on the solarpanel.

To realize the three solar panel configurations, the solar antenna has to sharecopper pads with the alternative large solar cell. If used for the solar antenna,the pads must not be connected to electrical ground. If the second largesolar cell is applied, the shared pads need to be connected to the other solarantenna pads. The allocation of copper pads to for the antenna and the solarcells is shown in figure 6.4. A 0 W bridge is provided on the back of the solarpanel, that allows to connect the solar antenna pads with the solar cell padsin case a large solar cell is used (cf. figure 6.6(a)).

6.2.2 Power Generation and Distribution

Power is generated either by two large solar cells, or a combination of one largeand four small solar cells together with the cell on top of the solar antenna.As described in the previous section, the solar antenna and solar cell footprintpads on the panel allow for connecting the two different setups already. Inorder to have a footprint design, that in addition supports the use of largeand small cells in the same place, a gridded design is applied to all solar cellfootprint pads (cf. figure 6.4 and figure 6.6(b)).

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Figure 6.4: Allocation of antenna and solar cell pads

Solar cell power is collected and combined using a set of bypass diodes onthe back of the panel. From this point, electric power is guided through theMPPT hardware, and then further towards the single connector used for bothpower and data bus. MPPT uses a perturb and observe algorithm for findingthe maximum power point, which depends on multiple factors like solar celltemperature and the connected load (cf. [125]).

Supply power is not generated locally on the solar panel. Instead, it isdrawn from the power bus lines, which are supposed to be controlled by thesuperordinate PCDU node of the satellite. This node monitors the powerdrawn by the panel and latches it off in case of excessive power consumption.Current-sense amplifiers and power-distribution switches located on the solarpanel allow to locally monitor and latch subsystems like the sun sensor, themagnetic actuator, or the pFDA (cf. Grau, Tschoban, et al. in [65]). Therefore,malfunction of one of the local subsystems does not jeopardize the completesolar panel functionality or the satellite bus.

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6.2.3 Attitude Determination

For attitude determination one TU Berlin sun sensor plus two MEMS magneticfield and two angular rate sensors are utilized on the panel. On the BEESATsatellites, the sun sensor is already located on the solar panels. It consists of aPSD, aperture, and amplifier circuitry [126]. Amplified signals are processedon the OBC board, and there combined with magnetic field and angularrate sensor data to estimate the attitude. In the present design, having allsensors and a microcontroller on the solar panel allows to perform attitudedetermination locally, and provide attitude estimates from each solar panel tothe ADCS core, that is running on the OBC.

Implementation of attitude determination hardware and algorithms is designedbased on BEESAT heritage. This reduces the likeliness of failure in case themulti-functional solar panel will be used on future CubeSat missions.

6.2.4 Attitude Actuators

Two types of attitude actuators are used on the solar panel: an optimizedmagnetic actuator for detumbling and coarse pointing (cf. section 4.4.3)together with a panel-mount pFDA for precise and agile maneuvers (cf.section 5.4.6).

6.2.4.1 Magnetic Actuator Design

The magnetic coil is realized on eight internal layers of the PCB. The coil shouldproduce a minimum magnetic dipole of 45 mA m2, and consume a maximumof 200 mW of electric power at a supply voltage of 3.30 V. Track separationis set to 100 µm, and the values for all other technological parameters areset to the required minimum of the board manufacturer. Using the magneticactuator optimization procedure (cf. section 4.3), a number of windings perlayer of 19 and a track width of 550 µm is found. GNU Octave is used todisplay (cf. figure 6.5) and export the coil geometry to KiCAD, the freeelectronic design automation (EDA) software used to create all schematics andPCB layouts for the solar panel and the picosatellite fluid-dynamic actuator.

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-40 -30 -20 -10 0 10 20 30 40-50

-40

-30

-20

-10

0

10

20

30

40

50

x [mm]

y[m

m]

Figure 6.5: Magnetic coil layout

The magnetic coil driver is implemented using an H-bridge, which is controlledby the attitude control algorithms executed on the microcontroller on thesolar panel.

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6.2.4.2 Panel-Mount Pico Fluid-Dynamic Actuator

As precise and agile attitude control actuator, the pFDA shown in figure 5.10on page 76 is integrated with the solar panel. Solar panel design work withrespect to the actuator comes down to the mechanical and electrical interface,including power monitoring and latching.

Monitoring and latching circuitry is already discussed in section 6.2.2. Asthe pFDA implements an external I2C data interface, the physical electricalinterface is realized using board-to-board connectors between the solar paneland pFDA PCBs. The interface provides logic and pump power to the actuator,and uses one of the two I2C buses that are available on the solar panel forcommunication.

6.2.5 Command and Data Handling

A 32-bit, low-power microcontroller is used as central processing unit on thesolar panel. The controller connects to the redundant CAN bus using twotransceiver chips. Algorithms for MPPT and attitude estimation are executedthere, and the determined attitude is provided to the data bus. Commandsfor the attitude actuators are received and either translated and passed on tothe pFDA, or directly executed for the coil driver. The implemented PCDUsoftware monitors the subsystem power demands, and latches them in case ofexcess.

6.3 Solar Panel Assembly and Test

Assembly of the multi-functional solar panel was carried out in multiple stepsbetween Fraunhofer IZM and the author. Integration of the solar antenna,electric component placement, and solar cell integration were carried outat Fraunhofer IZM. Pre-assembled panels were then shipped to TU Berlin,integrated with the pFDA, and the microcontroller programmed with thesoftware. Thermal test were then conducted at TU Berlin, before the solarpanels were shipped back to Fraunhofer IZM for mechanical tests.

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6.3 Solar Panel Assembly and Test 111

6.3.1 Assembly

Images of the front and the back of the pre-assembled solar panel are shown infigure 6.6. The back shows the soldermask islands that are required for gluingthe PCB substrate directly to the pFDA or the primary satellite structure,respectively (cf. figure 6.6(a)). This image also shows the pFDA outline,where no electronic components can be placed, and the opening required forthe antenna connector. At this step, the front of the solar panel does nothold any components, which allows to see the gridded structure of the solarcell and antenna pads in figure 6.6(b).

(a) Back (b) Front

Figure 6.6: Pre-assembled multi-functional solar panel

Figure 6.7 shows images of the front and the back of the fully assembledmulti-functional solar panel. The solar antenna was integrated at FraunhoferIZM, who also equipped the solar panel with electronic components. The solarcells were glued to the panel and the pads of the smaller cells wire-bonded tothe pads on the PCB. The actuator was fully assembled by the author, andthen integrated with the pre-assembled solar panel.

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(a) Back with actuator (b) Front with solar antenna

Figure 6.7: Fully assembled multi-functional solar panel

After fully assembling the solar panel, a fit check using a 3D-printed mock-upof the adapted BEESAT satellite structure was performed (cf. figure 6.8). Nodifficulties were encountered during fit-check, and the solar panel was easilyassembled and disassembled.

Finally, the solar panel mass was determined for the three different versions.The basic version weighs approximately 40 g, which is very competitive, ifcompared against the solar panels listed in tables 6.1 and 6.2, as there are onlytwo other solar panels, that weigh less. This seems quite interesting, as thepresented multi-functional solar panel includes more electronic componentsthan any of the commercially available panels.

The mass of the enhanced attitude control version, which includes the pFDAmass (cf. chapter 5), was determined to approximately 85.5 g. This versionweighs about twice of the average solar panel available on the market, butoffers the unique feature of having two different attitude control actuatorsintegrated on the same panel.

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(a) Adapted structure (b) Solar panel on adapted structure

Figure 6.8: Fit-check of the fully assembled multi-functional solar panel

With the additional solar antenna integrated (cf. page 35), the mass of thecommunication and enhanced control version is approximately 104 g. Thefully equipped solar panel has a mass that is roughly equivalent to 7.78 % of astandard single unit CubeSat. While the advantage of using highly integratedmulti-functional solar panels is not obvious on first sight, it is explained indetail in section 6.4.

6.3.2 Functional Verification

Functional verification carried out for some of the components found onthe multi-functional solar panel has been addressed in earlier sections (cf.sections 3.4 and 5.6). The sun sensor [126] or the magnetic field and angularrate sensors have been functionally verified in other projects at TU Berlin, andhave gained flight heritage on multiple spacecraft in orbit. MPPT hardwareand software was functionally verified at Fraunhofer IZM using a demonstratorwith one large and five small solar cells and a development board of themicrocontroller which is used on the solar panel [125]. The only component

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that requires functional verification on system level is the optimized magneticactuator.

For functional verification of the embedded air coil, power consumption andmagnetic dipole were measured. At the targeted supply voltage of 3.30 V, thecoil consumes 205 mW of power. This is only a 2.30 % difference from thetargeted value (cf. section 6.2.4.1). In order determine the magnetic dipole𝜇, the magnetic flux density 𝐵 was measured at a defined distance from thesolar panel along a line that passes perpendicularly through the center of thecoil. Distance 𝑥 was chosen to be large enough, so that the relation betweenmagnetic dipole 𝜇 and magnetic field strength 𝐵 is given by

𝐵 = 𝜇04𝜋

2𝜇

𝑥3 . (6.1)

Here, 𝜇0 is vacuum permeability. From this equation and the sensor data,magnetic dipole was calculated to be 47.5 mA m2. This value is differing fromthe targeted value by 5.56 %.

6.3.3 Environmental Tests

Environmental tests were carried out by a colleague of the author, and onlypreliminary results were published in [64]. Vacuum tests still need to beconducted. Tailored thermal tests were carried out at TU Berlin accordingto ECSS with a reduced number of cycles. Thermal test results showed nofailure of components.

For mechanical tests, two fully integrated solar panels were mounted in acarrier (cf. figure 6.9), and tested on a shaker at Fraunhofer IZM. Duringvibration tests, the solar antennas teared off from their mounting points, andthe solar cells on the antenna and around the antenna got partially destroyed.Analysis of the pads on the solar panel and the bottom of the antenna shows,that not enough adhesive was applied during integration of the antenna withthe panel at IZM (cf. figure 6.10).

Except for the torn-off solar antenna, no other subsystem showed signsof damage after mechanical testing. Fluid conduits didn’t show any signsof leakage, and all subsystems operated nominal. Due to the fact, that

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Figure 6.9: Solar panels mounted for mechanical tests

(a) Back (b) Front

Figure 6.10: Damaged solar panel and torn-off solar antenna

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the solar panels are connected to the frame using screws and adhesive tosimulate the assembly situation found on the satellite, the panels can’t beremoved from the frame. Therefore, no repetitive functional verification of theactuators’ dynamical properties was possible. Measurements of actuator powerconsumption, however, showed results comparable to the ones presented insection 5.6.1. It is therefore assumed that neither pump driver electronics norpump components have suffered any damage from mechanical testing.

6.4 Highly Integrated Multi-Functional Solar PanelAdvantages

Following the successful development of a multi-functional solar panel andall of its components in the previous sections, the realization of a denselyintegrated CubeSat with a large and compact payload volume is close athand. To better understand the implications of such panels for CubeSat

Table 6.3: Components of a proposed high-density single unit CubeSat design

name qty masssingle total

g gOBC 1 70 70PCDU 1 70 70battery compartment 1 174 174communication panel 1 103.5 103.5control panel 2 85.5 171basic panel 3 40 120shell 1 89 89lid 2 34 68tertiary structure 1 45 45harness 1 55 55bus 965.5payload 364.5

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6.4 Highly Integrated Multi-Functional Solar Panel Advantages 117

design, an updated version of the rough satellite model presented in figure 6.1was created. This design is based on the assumptions for the integrationdensity of the required subsystems on page 98. Components of the proposeddesign are listed in table 6.3. Besides the names and quantities, the table alsoholds component mass and calculates the totals. Maximum payload mass isestimated under the assumption, that total satellite mass is limited to 1.33 kg.

Mass could be measured for the three solar panel configurations (cf. page 112).Masses of the other components are based on values presented in [54]. OBCand PCDU masses are increased to account for the additional UHF transceivertechnology embedded there, and the assumed increased component countdue to denser integration. Mass of the shell and the two lids was adapted,to account for the changes required for actuator mounting. Masses given forthose main structural parts are calculated from CAD. Battery compartment,tertiary structure, and harness mass were taken directly from the referenceddocument. This leads to a total payload mass of 365 g, which representsabout 27.4 % of the total mass. Applying a total mass limit of 2 kg, availablepayload mass ratio would be about 1 030 g, which is equivalent to 51.7 %.

Comparison with the stated design criteria of 1 kg payload mass on a 2 kgsingle unit CubeSat derived in section 2.2.1 shows the good fulfillment usingthe highly integrated, multi-functional solar panel.

Due to the complex geometric boundaries of the pFDAs on the panels, availablepayload volume is not estimated based on tables and calculations. The CADmodel of the satellite shown in figure 6.11 is used for this. The subsystemsare colored similarly to figure 6.1. The payload, shown in yellow here, hasa volume of 558 cm3. Under the assumption, that the maximum volume ofa single unit CubeSat is 1 000 cm3, this is 55.8 % of the total volume. Likeavailable payload mass, available payload volume is meeting the 50 % designcriteria derived in section 2.2.1 for single unit CubeSats.

Section 2.2.1 also states a requirement of 2 W of average payload power.Without well-defined mission objective and payload requirements, estimationof available power is difficult. Like for mass estimation, looking at BEESAT-4 provides an insight to this. Power consumption estimated in [54] givesan average stand-by power consumption of about 500 mW. For a typicaloperations scenario, the same source states a power consumption of about

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Figure 6.11: CAD model of a high performance single unit CubeSat

750 mW. The scenario is tailored to the payload of BEESAT-4, and doesnot require active attitude control over longer periods of time. If a highperformance payload, that utilizes the additional payload resources in terms ofmass and volume, has increased pointing demands over longer periods of time,the average power directly available for the payload is drastically reduced.Even taking the reduced power demand of the pFDAs (cf. section 5.6.1) andthe optimized magnetic actuator into account, this might require to equip thesatellite with additional, deployable solar panels.

Fulfillment of the other design criteria derived in sections 2.2.2 to 2.2.4 wasalready demonstrated in the previous chapters of this thesis. In conclusion,the utilization of the presented highly integrated, multi-functional solar paneltogether with miniaturized subsystems allows to bring the integration densityof single unit CubeSats to a level similar to that of the larger triple unitCubeSats.

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7 Summary and Conclusion

The objective of this thesis is to investigate potential technologies to increasethe integration density of CubeSats and apply them for the design andimplementation of a multi-functional CubeSat solar panel.

7.1 Summary

Research in this works starts with a brief analysis of the evolution of CubeSatlaunches over the last fifteen years (cf. chapter 1). During the first decade,university satellite missions were foremost used for space engineering education.The prevalent form factors during that time were single and double unitCubeSats. Growing acceptance of the CubeSat standard, first published in1999, lead to a wider acceptance of CubeSat missions not only for educationalpurposes, but as the basis for market activities of the so-called New Spacecompanies. Since 2013, launch numbers are growing at a large pace andhigh-performance triple unit CubeSats are now dominating the market. Non-commercial missions are showing a smaller increase in launch numbers. A closeanalysis reveals that the majority of CubeSats launched by universities are stilltechnology demonstration missions using single to triple unit spacecraft. Thisis slowly shifting to larger spacecraft used for space science, demonstrated forexample by the launch of the QB50 constellation in 2017.

Observations in chapter 2 show, that the growing number of stakeholders in theCubeSat market may be sorted into three categories: CubeSats utilized in large-scale, single purpose constellations are developed by independent commercialactors. These are contrasted by commercially-procured spacecraft, that areordered by stakeholders interested in the utilization of payloads. Finally,independent university CubeSats are developed by flagship universities andconstantly push the technological state of the art [22]. In-depth analysis of

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120 7 Summary and Conclusion

integrated CubeSat platforms for commercially procured satellites in this workreveals, that the usage of standardized components and subsystems enableshigh performance multi-purpose, multi-user buses for triple unit CubeSats.Implementation of high performance payloads on single unit CubeSats is,however, hindered by standardized subsystems due to their poor volumetricoptimization.

The insight resulting from the authors observation of the CubeSat marketis used to derive design criteria for high performance single unit CubeSats(cf. section 2.2). A core requirement is the increase of available payloadvolume, mass, and power budgets together with higher data throughput. Apromising approach to enhance payload capability of single unit CubeSatsis the relocation of multiple different functionalities from inside the stack ofCubeSat subsystems to the solar panels [63]. Based on the aforementionedcriteria for the complete satellite, design criteria for highly integrated, multi-functional solar panels are derived (cf. section 2.3). Three major possibilitiesto increase the integration density are identified: First the development of asolar patch antenna for high-throughput payload data downlink in the S-band.Second the optimization of magnetic actuators. And third the replacement ofminiaturized reaction wheels with a novel class of attitude control actuators,so-called pFDAs.

Research and development according to these design criteria was conductedin a joint project between researchers at Fraunhofer IZM and the Chair ofSpace Technology at TU Berlin under the lead of the author. Investigation ofsolar patch antennas was carried out by Fraunhofer IZM, and their results aresummarized in chapter 3.

Analysis of magnetic attitude control actuators reveals poor optimization withrespect to magnetic dipole, power consumption, and mass (cf chapter 4). Anovel model-based optimization approach for magnetorquer design is formu-lated by the author [75]. The approach allows to define a set of equality andinequality constraints for aforementioned magnetorquer characteristics. Butalso for additional geometric dimensions and technology parameters. Resultsfrom the application of the optimization procedure reveal a big optimizationpotential for magnetic actuators, especially for embedded air coils.

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7.2 Conclusion 121

Recent advancements in the development of fluid-dynamic actuators leadto the proposition of picosatellite fluid-dynamic actuators with rectangular,3D-printed conduits for CubeSat attitude control (cf. chapter 5, [64, 109]).Calculation of angular momentum for flat and three-dimensional conduits isderived, and flat square conduits are identified to be best suited for CubeSatapplications due to their better volumetric utilization. Functional verificationis carried out for those conduits and miniaturized pump driver electronics,and an automated model-based parameter estimation is applied to identifythe dynamical properties of the actuators. Identified pFDA parameters areused for a comparison of the agility of miniaturized reaction wheels and fluidactuators. Panel-mount pFDAs offer no possibility for redundant assemblies.It can be shown that proper design of three-dimensional actuators enables theuse of pFDAs in redundant configurations (cf. section 5.8).

Work in this thesis culminates in the design and implementation of a highlyintegrated, multi-functional solar panel. Based on the characteristics deter-mined for the fully assembled solar panel, a CubeSat design is presented thatis in accordance with the design criteria derived in section 2.2.

7.2 Conclusion

Research conducted by the authors shows, that the integration density onsingle unit CubeSats can be increased by utilizing highly integrated, multi-functional solar panels. In order to achieve higher levels of integration on thepanels as well as in the complete satellite, optimization and miniaturization ofcomponents need to be applied consistently. The resulting integration densityleads to a significant increase in available payload resources.

One component that showed big potential for optimization are magneticattitude control actuators. The optimization procedure developed by the authorin the scope of this work allows to improve the characteristics of magnetorquers,and here especially of embedded air coils. This was demonstrated by thedesign of a coil for the integrated solar panel that produces the same amountof magnetic dipole at two thirds the power consumption of the coil which iscurrently used on TU Berlin’s CubeSat missions.

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122 7 Summary and Conclusion

Investigation of picosatellite fluid-dynamic actuator miniaturization carried outby the author led to the result that this type of actuator is ready for utilizationon CubeSat missions. Its application on the highly integrated, multi-functionalsolar panel contributes most to the increase seen for available payload volumein the proposed high-density CubeSat design. The level of miniaturizationachieved is possible due to modern additive manufacturing technologies andthe availability of miniaturized electronic components.

Research conducted by the author on three-dimensional picosatellite fluid-dynamic actuator conduit geometries revealed the feasibility of redundantactuator concepts for single unit CubeSats. The developed conduit geometryallows for dense integration of the actuators with satellite subsystems andsecondary structure, which further enhances CubeSat capabilities. Individualthree-dimensional actuators have been functionally verified in the lab, whilethe full assembly awaits verification on a sounding rocket experiment scheduledto launch in spring 2019.

Consequent application of the methods and technologies developed by theauthor for future CubeSat designs will lead to a significant increase of payloadcapabilities using the highly integrated, multi-functional solar panel. If redun-dant actuator concepts are required for such missions, the multi-functionalpanels can be combined with three-dimensional picosatellite fluid-dynamicactuator designs.

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A Magnetic Coil Data

Table A.1: Wound torque rod data

id 𝜇 𝑃 𝑀 𝑈 refmA m2 mW g V

32 200 200 30 5 [127]33 200 300 7 5 [96]34 100 295 3 5 [96]35 200 140 9 5 [96]39 240 188 27.5 2.50 [128]41 200 130 22 5 [129]42 240 208 22 2.50 [130]43 209 46 31.5 4 [78]54 217 200 30 8.20 [79]55 193 150 50 8.20 [79]56 107 100 49 8.20 [79]57 106 298 32 8.20 [79]58 70.7 200 18.2 8.20 [79]79 157 322 29 3.30 [90]a 113 288 2.94 5 [75]b 101 247 2.88 5 [75]

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138 A Magnetic Coil Data

Table A.2: Embedded air coil data


21 44 285 3.60 5 [77]22 39.5 322 3.20 5 [77]23 66.4 293 8.20 5 [82]24 42.9 298 5.90 5 [82]28 50 250 6 1.60 [83]49 38 81 6 3.30 [80]50 131 259 6 3.30 [81]g 67.1 300 5.93 3.80 [75]h 40 105 5.98 1.80 [75]i 40 291 1.84 3.30 [75]j 40.4 285 1.89 1.80 [75]k 40.2 288 2.10 1.80 [75]

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139

Table A.3: Wound air coil data


1 28 81 5.80 3.80 [87]5 105 228 19.2 4.50 [89]6 80.6 174 19.2 4.50 [89]7 67.5 146 19.2 4.50 [89]8 58 125 19.2 4.50 [89]9 35.3 50 11.7 3.30 [131]

10 45 50 20 3 [85]12 100 260 31.2 5 [132]13 100 140 50.3 5 [132]14 97.3 226 19 2.40 [133]16 80 155 9.90 3.30 [90]17 88 245 9.70 3.30 [90]18 101 350 10.2 3.30 [90]19 139 74 106 3.30 [86]48 100 100 40 NaN [134]51 111 78 27 3.60 [135]52 106 333 21 3.70 [84]59 83.1 179 13.5 3.30 [90]61 93 221 14.4 3.30 [90]63 103 273 15.4 3.30 [90]65 114 332 16.6 3.30 [90]87 83.5 179 21.9 3.30 [90]88 86.3 299 29.9 3.30 [90]89 93 221 24.8 3.30 [90]90 95 299 29.9 3.30 [90]91 103 273 28.2 3.30 [90]92 104 298 29.9 3.30 [90]d 30.1 83 5.82 7.60 [75]e 27.6 70 5.83 3.80 [75]f 27.6 83 4.89 3.80 [75]

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Schriftenreihe Institute of Aeronautics and Astronautics: Scientific SeriesHrsg.: Prof. Dr.-Ing. Dieter Peitsch, Prof. Dr.-Ing. Andreas Bardenhagen,Prof. Dr.-Ing. Klaus Brieß, Prof. Dr.-Ing. Robert Luckner,Prof. Dr.-Ing. Julien Weiss

ISSN 2512-5141 (print)ISSN 2512-515X (online)

01: Behrend, Ferdinand: AdvancedApproach Light System. Der Einfluss eineszusätzlichen visuellen Assistenzsystems zurSteigerung des Situationsbewusstseins beikritischen Wetterbedingungen hinsichtlichvertikaler Fehler im Endanflug. - 2017. -XXIV, 206 S.ISBN 978-3-7983-2904-1 (print) EUR 16,50ISBN 978-3-7983-2905-8 (online)DOI 10.14279/depositonce-5819

02: Gordon, Karsten: A flexible attitudecontrol system for three-axis stabilizednanosatellites. - 2018. - XXVI, 173 S.ISBN 978-3-7983-2968-3 (print) EUR 14,00ISBN 978-3-7983-2969-0 (online)DOI 10.14279/depositonce-6415



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ISBN 978-3-7983-3026-9 (print)ISBN 978-3-7983-3027-6 (online)

Institute of Aeronautics and Astronautics: Scientific Series Band 3

This thesis investigates potential technologies to increase the integration density of CubeSats. Observations of the CubeSat market and missions are recorded in order to derive design crite-ria for high performance single unit CubeSats. A promising approach to increased integration density is relocation of the components of multiple satellite subsystems to form a highly integ-rated, multi-functional solar panel. Eligible components are usually allocated to the communi-cation system, the electric power system, or the attitude determination and control system. In a joint research project, development, optimization, and miniaturization of those components in order to form a highly integrated, multi-functional solar panel is investigated.The author first summarizes the development work of the project partners for a picosatellite solar antenna and puts it into relation to the overall solar panel design. Advantage of using solar antennas over simple patch antennas is the reduced loss of solar cell area, and hence available electric power, that is usually accompanied by the usage of higher frequency bands for broadband payload data transmission.

Se

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9 783798 330269I S B N 9 7 8 - 3 - 7 9 8 3 - 3 0 2 6 - 9 http://verlag.tu-berlin.de

Sebastian Grau


contributions to the advance of the integration density of cubesats · instute of aeronaucs and...

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