Sebastian Grau
Contributions to the Advance of the Integration Density
of CubeSats
Sebastian Grau
Contributions to the Advance of the Integration Density of CubeSats
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
Institute of Aeronautics and Astronautics: Scientific Series | 3
Sebastian Grau
Contributions to the Advance of the
Integration Density of CubeSats
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Gutachter: Prof. Dr.-Ing. Hakan Kayal (Universität Würzburg) Die Arbeit wurde am 30. August 2018 an der Fakultät V unter Vorsitz 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 4.0 http://creativecommons.org/licenses/by/4.0/
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Print: Pro BUSINESS
Layout/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 Repository of the Technische Universität Berlin
DOI 10.14279/depositonce-7293
Contents
1 Introduction 1
1.1 Recent and Future Evolution of CubeSat Launches . . . 2
1.1.1 NewSpace Constellations . . . 2
1.1.2 Civil Constellations . . . 3
1.1.3 University-Class Spacecraft . . . 4
1.2 Thesis Outline . . . 5
2 High Performance Single Unit CubeSat Design Approaches 9 2.1 CubeSat Market Observations . . . 9
2.1.1 Independent Commercial Spacecraft . . . 10
2.1.2 Commercially-Procured Spacecraft . . . 11
2.1.3 Independent University Spacecraft . . . 17
2.2 High Performance Single Unit CubeSat Design Criteria . . . 23
2.2.1 Provided Payload Resources . . . 23
2.2.2 Attitude Determination and Control Capabilities . . . 23
2.2.3 Downlink Capabilities . . . 24
2.2.4 Data and Power Bus . . . 24
2.3 Highly Integrated Solar Panel Design Criteria . . . 25
3 Picosatellite Solar Antenna 27 3.1 Solar Antenna State of the Art . . . 27
3.2 Patch Antenna Development . . . 29
3.3 Solar Cell Integration . . . 33
3.4 Solar Antenna Functional Verification . . . 35
4 Magnetic Actuator Optimization 37 4.1 Magnetic Actuator State of the Art . . . 38
4.1.1 Wound Torque Rods . . . 39
4.1.2 Embedded Air Coils . . . 39
4.2 Magnetorquer Optimization State of the Art . . . 42
4.2.1 Experiment-Based Optimization . . . 43
4.2.2 Model-Driven Optimization . . . 43
4.2.3 Sequential Quadratic Programming . . . 44
4.3 Formulation of a Novel Optimization Procedure . . . 45
4.3.1 Mathematical Modeling of Magnetic Coil Properties . 46 4.3.2 High-Dimensional Parameter Space Implementation . 47 4.3.3 Discrete and Integer Input Parameters . . . 48
4.3.4 Flexible Optimization Objectives Definition . . . 48
4.3.5 Global Optimum Search . . . 50
4.4 Optimization Results . . . 52
4.4.1 Wound Torque Rod Optimization Results . . . 52
4.4.2 Wound Air Coil Optimization Results . . . 52
4.4.3 Embedded Air Coil Optimization Results . . . 54
5 A New Class of Picosatellite Attitude Actuators 55 5.1 Technological Evolution of Fluid Spacecraft Actuators . . . . 55
5.2 Objectives for Actuator Miniaturization . . . 57
5.3 Fluid-Dynamic Actuator Fundamentals . . . 59
5.3.1 Planar Actuators . . . 60
5.3.2 Three-Dimensional Actuators . . . 63
5.3.3 General Spacecraft Dynamics . . . 67
5.4 Fluid Actuator Conduits for CubeSat Applications . . . 68
5.4.1 Conduit Considerations . . . 68
5.4.2 Pump Housing . . . 69
5.4.3 Actuator Electronics . . . 70
5.4.4 Manufacturing of Monolithic, Integrated Conduits . . 70
5.4.5 First Rapid Prototyping Experiences . . . 71
5.4.6 Conduit Geometries for CubeSat Applications . . . . 72
5.4.7 Advantages of Conduit Rapid Prototyping . . . 74
5.5 Driver Electronics . . . 74
5.5.1 Electronics Miniaturization . . . 74
5.5.2 Flexible Development Platform . . . 76
5.6 Functional Verification . . . 78
5.6.1 Power Consumption . . . 80
5.7 Fluid-Dynamic Actuators and Reaction Wheels . . . 85
5.7.1 Angular Rate and Acceleration . . . 86
5.7.2 Time-Optimal Slew Maneuvers . . . 86
5.7.3 Analysis of Traveled Angles . . . 87
5.8 Redundancy Concepts . . . 89
5.8.1 L-Shaped Conduits . . . 89
5.8.2 Crown-Shaped Conduits . . . 92
6 A Highly Integrated Single Unit CubeSat Solar Panel 95 6.1 Current Integration Density of CubeSat Solar Panels . . . . 98
6.1.1 Mechanical Properties . . . 99
6.1.2 Power Generation . . . 100
6.1.3 Attitude Determination Sensors . . . 101
6.1.4 Attitude Control Actuators . . . 102
6.1.5 Harness . . . 103
6.1.6 Conclusion . . . 104
6.2 Design of a Highly Integrated, Multi-Functional Solar Panel . 105 6.2.1 Communication . . . 106
6.2.2 Power Generation and Distribution . . . 106
6.2.3 Attitude Determination . . . 108
6.2.4 Attitude Actuators . . . 108
6.2.5 Command and Data Handling . . . 110
6.3 Solar Panel Assembly and Test . . . 110
6.3.1 Assembly . . . 111
6.3.2 Functional Verification . . . 113
6.3.3 Environmental Tests . . . 114
6.4 Highly Integrated Multi-Functional Solar Panel Advantages . 116 7 Summary and Conclusion 119 7.1 Summary . . . 119
7.2 Conclusion . . . 121
List of Figures
1.1 Number of CubeSat launches 2003–2022 . . . 1
1.2 Distribution of civil CubeSat launches 2003–2022 . . . 3
1.3 Distribution of university CubeSat application over form factor 2003–2018 . . . 5
3.1 Radiation mechanism of a patch antenna and integration of a solar cell . . . 29
3.2 Return loss and axial ratio of a quasi-square patch antenna for different thicknesses of the dielectric medium . . . 30
3.3 Return loss and axial ratio of all researched patch antennas . 31 3.4 Right hand and left hand circular polarization antenna gains for all researched patch geometries in the xz plane in dB . . 32
3.5 Right hand and left hand circular polarization antenna gains for all researched patch geometries in the yz plane in dB . . 32
3.6 Cross section schematic of solar antenna . . . 33
3.7 Image of the fully integrated solar patch antenna . . . 34
3.8 Solar antenna return loss measurement . . . 35
3.9 Solar antenna gain measurement . . . 36
4.1 Overview of magnetic actuator properties . . . 40
4.2 Overview of optimized magnetic actuator properties . . . 53
5.1 First demonstrator of a CubeSat fluid-dynamic actuator on an air bearing test facility . . . 58
5.2 Circular conduit . . . 61
5.3 Rectangular conduit . . . 63
5.4 L-shaped conduit . . . 64
5.5 Crown-shaped conduit . . . 66
5.6 First monolithic, integrated fluid-dynamic actuator conduit . 72 5.7 Monolithic, integrated fluid-dynamic actuator conduit . . . . 73
5.8 Circular integrated conduit used for student projects . . . 73
5.9 Block diagram of initial pFDA electronics . . . 75
5.10 First fully-integrated picosatellite fluid-dynamic actuator opti-mized for CubeSat application . . . 76
5.11 Components of the flexible development platform . . . 77
5.12 Flexible development platform assembled for air bearing use 77 5.13 First demonstrator of a pFDA using a 3D-printed conduit . . 79
5.14 Angular rate measurements of a pFDA . . . 79
5.15 Pump power demand at 5 V in mW . . . 81
5.16 Pump power demand at 12 V in mW . . . 81
5.17 Actuator parameter estimation based on a first-order LTI model 83 5.18 Estimated actuator parameters . . . 84
5.19 Comparison of simulated actuator properties . . . 87
5.20 Comparison of simulated time-optimal slew maneuvers . . . 88
5.21 Maximum angles traveled using time-optimal and non-optimal maneuvers . . . 89
5.22 L-shaped conduits in a redundant configuration . . . 90
5.23 Experimental assembly of redundant fluid-dynamic actuator configuration based on L-shaped conduits . . . 91
5.24 Crown-shaped conduits in a redundant configuration . . . . 93
6.1 Schematic view of a proposed high density single unit CubeSat using highly integrated, multi-functional solar panels . . . . 96
6.2 Nodes of a proposed high density single unit CubeSat . . . . 97
6.3 Solar cell orientation and mounting hole placement . . . 100
6.4 Allocation of antenna and solar cell pads . . . 107
6.5 Magnetic coil layout . . . 109
6.6 Pre-assembled multi-functional solar panel . . . 111
6.7 Fully assembled multi-functional solar panel . . . 112
6.8 Fit-check of the fully assembled multi-functional solar panel 113 6.9 Solar panels mounted for mechanical tests . . . 115
6.10 Damaged solar panel and torn-off solar antenna . . . 115
List of Tables
2.1 Integrated CubeSat platforms: mechanical and electrical
char-acteristics . . . 12
2.2 Integrated CubeSat platforms: attitude determination and control system characteristics . . . 16
2.3 Integrated CubeSat platforms: communication system charac-teristics . . . 16
5.1 Comparison of dynamical properties of CubeSat reaction wheels and the picosatellite fluid-dynamic actuator . . . 86
6.1 Single unit CubeSat solar panel properties . . . 99
6.2 Single unit CubeSat solar panel components . . . 102
6.3 Components of a proposed high-density single unit CubeSat design . . . 116
A.1 Wound torque rod data . . . 137
A.2 Embedded air coil data . . . 138
Glossary
AraMiS
Italian for Modular Architecture for Satellites; a project to research small satellite modularization, conducted at Politecnico Torino, Italy. 103, 105
Astrofein
Astro- und Feinwerktechnik Adlershof GmbH, a Berlin, Germany based manufacturer of satellite components. 86
BEESAT
Berlin Experimental and Educational SATellite (BEESAT), a series of CubeSats designed and operated by Technische Universität Berlin. Also the name of the first single unit CubeSat mission at TU Berlin, which was 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 demonstrate advanced attitude determination and control capabilities. 17, 18
BEESAT-4
The fourth in the series of BEESAT satellites, designed to demonstrate orbit determination using GPS and act as a target satellite for BIROS. 6, 17–19, 24, 25, 117, 118
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 container in 2019. 5, 6, 17, 21, 22, 24, 25, 98
BIROS
The Bi-spectral Infrared Optical System (BIROS) is successor to the TET-1 satellite and part of the DLR FireBIRD mission. It released BEESAT-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 CubeSat vendor. 11, 15
CanX-1
Canadian Advanced Nanospace eXperiment 1 (CanX-1) was launched as 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, in the scope of the AraMiS-C1 project. 105
Dove
The Dove spacecraft are triple unit CubeSats, built and operated by American 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 high efficiency solar cells. 28
GNU Octave
Free software that features a high-level programming language, primarily intended for numerical computations. 47, 48, 50, 75, 108
GomSpace
GomSpace is a Danish CubeSat and nanosatellite manufacturer based in Aalborg, Denmark. 11, 14
GOMX-3
In-orbit demonstration mission developed by GomSpace and funded by ESA. 11, 20
HISPICO
Highly Integrated S Band Transmitter for Pico and Nano Satellites devel-oped by Berlin based company iQ spacecom and Technische Universität Berlin. 24, 29
ISIS
Innovative Solutions In Space B.V. (ISIS) is a Dutch nanosatellite company located in Delft, The Netherlands. 11, 14
Kepler Communications
Kepler Communications Inc. is a Canadian satellite communication company 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 by American company Spire Global. 2, 11
MicroMAS
Micro-sized Microwave Atmospheric Satellite (MicroMAS) is a three-unit CubeSat developed by Massachusetts Institute of Technology Space Systems Laboratory. 4
MicroMAS 2
The Micro-sized Microwave Atmospheric Satellite 2 (MicroMAS 2) mission consists of two three-unit CubeSats developed by Massachusetts Institute of Technology Space Systems Laboratory and is the successor to the MicroMAS spacecraft. 4
MOVE-II
Munich Orbital Verification Experiment II (MOVE-II) is a single unit CubeSat developed at TU München and launched in 2018. 41
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 (OUFTI 1) 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 computer buses intended for extreme environment applications. 12, 15, 18–20, 23, 24, 95
PEEK
Polyether Ether Ketone, a high-temperature, high-strength thermoplastic polymer. 71
PID
Control law that uses a proportional, integral, and derivative component to 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 Miniaturized Components for Distributed Pico- and Nanosatellite Systems (BEESAT-5 to BEESAT-8). 5, 21, 23, 96
Planet Labs
American Earth observation company, developing and operating the Flock constellations formed by a large number of Dove satellites. 2, 10–12, 22, 95
Pumpkin
Pumpkin Space Systems, American CubeSat component supplier based in San Francisco, California. 11
PVC
Polyvinyl Chloride, plastic polymer used among others for electric cable isolation. 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) is a 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
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-band network of cooperating satellites, a non-CubeSat standard nanosatellite formation 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 for demonstration of the TUBiX20 satellite bus. 17, 57
TROPICS
NASA’s Time-Resolved Observations of Precipitation structure and storm 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 TUBiX20 satellite bus, successor to the TechnoSat mission. 17
TUBiX20
20 kg nanosatellite platform developed at TU Berlin, utilized for the missions TechnoSat and TUBIN. 24, 95–98
TUPEX-6
TU Berlin Picosatellite Experiment 6, a free-falling unit deployed from a REXUS sounding rocket to demonstrate redundancy concepts featuring fluid-dynamic actuators for picosatellites. 92
Tyvak
Tyvak Nano-Satellite Systems, Inc. is an American nanosatellite manu-facturer located in Irvine, California. 14
UHF
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 CubeSats offered 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
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 für
Luft- 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
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 Institut
für 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
M2M Machine-To-Machine. 2, 3
MEMS Micro-electro-mechanical System. 10, 17, 24, 70, 101, 108 MKI Multi-functional Integration of Miniaturized Satellite Components
for 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
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
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 m2s. 61–68, 84, 85, 90, 93
𝐻 Angular momentum vector in kg m2s. 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
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
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
𝜏𝑑 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
˙
𝜔 angular acceleration in °/s2. 82–84, 86 ˙
i i
i i
1 Introduction
In less than fifteen years, CubeSats developed from a tool in space engineering education 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 mission concepts using large numbers of CubeSats in constellations to gather data with high temporal and spatial resolution [5, 6]. This lead to a sharp increase in the number of CubeSat launches over the past years, which is predicted to continue (cf. [1, 2], figure 1.1). i i “tikzexternalize/fp7_cubesat_years” — 2019/5/23 — 23:24 — page 1 — #1 i i 2003 2004 2005 2006 2007 2008 2009 2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022 0 50 100 150 200 250 300 350 400 450 500 550 600 650 700 launch year numb er of Cub eSats launched failed announced predicted SpaceWorks forecast full market potential
1.1 Recent and Future Evolution of CubeSat Launches
Williams, Doncaster, and Shulman found that "overall sizes in the nanosatellite market are increasing to accommodate demand for additional payload capabil-ity", and that "the 3 U form factor is still expected to remain the standard in the 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 the single biggest contributor to nanosatellite launches [2]. The Dove satellites are triple unit CubeSats that provide complete coverage of the Earth with an optical resolution of 3–5 m and daily revisits [6]. Planet Labs alone is responsible for 35 % of all nanosatellite launches between 2008 and 2017 [2]. With Spire Global exists a second company that operates an established constellation of triple unit CubeSats [7]. Until February 2018, Spire Global successfully launched 67 of their Lemur-2 satellites. The forecast in [2] states the expectation, that Earth observation and remote sensing constellations will make up 50 % of the nanosatellite market, and communications constellations will account for 20 % over the next five years.
Among the new communication constellations is UK-based company Sky and Space who launched an inter-satellite link and communications proof-of-concept mission consisting of three triple unit CubeSats, called Three Diamonds, 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 deployed and operational by 2020" [10]. The so-called Pearl spacecrafts will form a mobile telecommunication network for the equatorial countries.
In January 2018, Canadian company Kepler Communications launched their first triple unit CubeSat as a demonstrator for their planned Ku-band
commu-i i
i i nications constellation [11]. The complete constellation will consist of 75–140
operational spacecraft [12].
Williams et al. identify further companies, that are developing communications constellations "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 2018 or 2019 [1, 2].
1.1.2 Civil Constellations
With regard to civil CubeSat missions, Williams et al. state that "despite rapid commercial market share growth, civil1(...) operator demand is expected to
remain consistent over the next 5 years" [2]. This statement is contrasted by data on civil launches taken from [1] and shown in figure 1.22: The announced
CubeSat launches for 2018 triple the number of launches in 2017.
1Williams et al. define civil operators as "operators whose primary satellite purpose is
non-military or non-profit activities" [2].
2The eight operator classes found in [1] have been regrouped according to the three
operator classes found in [2].
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Among those announced missions is NASA’s TROPICS mission; a scientific constellation of six identical triple unit CubeSats, which is currently under development and "on-track to deliver fligh-ready hardware in 2019", as stated by Blackwell, Burianek, Clark, et al. in [16]. The spacecraft is based on the MicroMAS [17] and MicroMAS 2 [16] missions, launched in July 2014 and January 2018, respectively. All satellites feature a commercially-procured bus, and a single unit payload for observation of precipitation structure and storm intensity of tropical cyclones [16].
The largest constellation of civil operated CubeSats is QB50 [5], consisting of a total of 36 double and triple unit spacecraft launched in two batches in April 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. In total, 13 ion-neutral mass spectrometers (INMSs) provided by MSSL, 19 flux-Φ-probe experiment (FIPEX) sensors provided by TU Dresden, and 11 multi-needle Langmuir probes (m-NLPs) supplied by UiO are located on the spacecraft of the constellation and will be used to study the properties of the upper thermosphere [20]. Each of the science units was developed to "half a CubeSat 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 that fly reliable and significant missions every few years, independent universities that have developed their own string of successful missions, and hobbyists with 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 of small satellite technology.
The first satellite constellation of TU Berlin, called S-Net, was successfully brought into orbit on February 1, 2018 [23]. The four satellites have dimensions of 25 × 25 × 25 cm [24] and are therefore not compliant with the CubeSat
i i
i i standard [25]. Primary mission objective is the demonstration of
multi-point inter-satellite link in the S-band between the members of a swarm of nanosatellites without active orbit control [24].
TU Berlin’s second constellation, the BEESAT-5 to BEESAT-8 satellites of the PiNaSys II mission, are scheduled for launch in 2018. The four satellites are quarter-unit CubeSats intended for demonstration of "a newly developed communications subsystem in the UHF band and an experimental GPS receiver", as stated by Baumann et al. in [26]. Beyond that, the satellites are designed almost complete single-fault tolerant and feature multi-functional optical attitude determination sensors [27]. Another specialty of the mission is, that the four satellites will be launched as a pack from one single unit CubeSat deployer [28].
1.2 Thesis Outline
In conclusion, the economic, civil, and university sectors show a clear trend towards more CubeSat missions to be launched and an increase in payload performance. University missions are starting to put more attention on scientific CubeSat missions (cf. [5], figure 1.3) At the same time, ongoing i
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i i
0.25U 1U 1.5U 2U 3U 6U 12U
0 15 30 45 60 75 90 105 120 form factor numb er of Cub eSat launches Space Activity Space Science Space Technology
Figure 1.3:Distribution of university CubeSat application over form factor 2003–
research on miniaturization of CubeSat subsystems allows to develop smaller buses. Combining these factors, the question arises to what extent it is possible to 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 of CubeSats. Research presented in this work was carried out in the scope of a joint, DFG-funded project titled Multi-functional Integration of Miniaturized Satellite Components for Increased Payload Capacity of Pico-Satellites (MKI) at the Chair of Space Technology at TU Berlin and Fraunhofer Institute for Reliability and Microintegration (IZM). Authorship of original work of the author or the project partners is highlighted in the following outline.
In chapter 2, the author discusses criteria and drivers for high performance CubeSat design. Three different CubeSat categories are identified from market observations, and their design characteristics are compared against each other. In this context, special attention is paid to missions of the two major German contributors to CubeSat development, launch, and operation: TU Berlin and Universität Würzburg. Analysis of the BEESAT-4, BEESAT-5 to BEESAT-8, and UWE-3 missions together with the results of the market observations are used to derive design criteria for high performance single unit Cubesats. One approach to meet those criteria is the use of highly integrated, multi-functional solar panels, which host components from different satellite subsystems. The author ends this chapter with the definition of design criteria for multi-functional solar panels.
To show, how formulated design criteria are met, each of the following chapters represents stand-alone research and development of subsystem components that reside on the multi-functional solar panel. Each chapter begins with a 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 communication from space to the ground. Work was carried out at Fraunhofer IZM, and supported by the author in an advisory capacity.
The author dedicates chapter 4 to numerical optimization of magnetic actuators for CubeSat applications. After outlining the state of the art of both actuators as well as their optimization, the formulation of a novel magnetic actuator optimization procedure is presented. Conclusively, results from the application of the proposed procedure are discussed.
Chapter 5 addresses the development of a novel CubeSat attitude control actuator, the so-called picosatellite fluid-dynamic actuator (pFDA), by the author. Technological evolution of fluid actuators is revised first, followed by the derivation of design objectives for fluid actuator miniaturization. To prepare the selection of the optimal fluid conduit geometry, angular momentum calculation is derived for flat and three-dimensional conduits. Following a section dedicated to conduit design and manufacturing, miniaturization of the pump driver is addressed. Functional verification results in the estimation of actuator dynamical properties. Based on the identified properties, a comparison of reaction wheels and pFDAs is carried out in order to asses the agility of the two fundamentally different actuators. The final section is dedicated to redundancy 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 to research the implementation of a highly integrated, multi-functional solar panel. Starting out from the analysis and discussion of the current integration density of CubeSat solar panels, a multi-functional solar panel design is addressed. Subsequently, assembly and test results are detailed before the chapter ends with the comparison between the realized solar panel and the criteria defined in section 2.3.
Finally, the author summarizes and concludes this thesis in chapter 7, and gives a recommendation for future work in the field of densely integrated CubeSat components and the application of pFDAs.
2 High Performance Single Unit CubeSat
Design Approaches
Satellites in the CubeSat constellations addressed in section 1.1 rely on inexpensive commercial off-the-shelf (COTS) components, which became available in recent years due to the technological advancement largely driven by the consumer electronics market, as stated by Bandyopadhyay in [29]. While Shimmin et al. in [30] state, that CubeSat subsystems built from COTS components 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 mass and volume, as well as electric power and link budget. In comparison to triple unit satellites, smaller CubeSats, especially below 2 U, were found to be rarely used for "science missions with relatively high performance", as stated by Poghosyan and Golkar in [4].
2.1 CubeSat Market Observations
While conforming to the CubeSat design specification (CDS) [25], design and composition of present-day CubeSats greatly varies depending on the developing/operating organization and their approach to satellite development. From the range of spacecraft addressed in section 1.1, three broader categories of CubeSats are derived:
Independent commercial CubeSats are completely designed and
manufac-tured in-house by expert companies to meet the requirements of a single constellation (cf. section 2.1.1).
Commercially-procured CubeSats are used by organizations, that are more
(cf. section 2.1.2). Therefore, those organizations procure the complete bus or the majority of the subsystems from CubeSat vendors.
Independent university CubeSats are developed by flagship or independent
universities1, as defined by Swartwout and Jayne in [22] (cf.
sec-tion 2.1.3). Those universities develop the complete satellite, and may also be responsible for the primary payload. This does not necessarily mean, that no parts of their CubeSats are procured, but they are able to avoid parts that have negative influence to volumetric utilization or power consumption on the spacecraft.
2.1.1 Independent Commercial Spacecraft
Probably the best example for independent commercial CubeSats are the Dove satellites of Planet Labs’ Flock constellation (cf. page 2). They feature a telescope of about 2.5 U volume on a triple unit spacecraft. Due to the telescope’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 center axis of the satellite connects telescope and detector. Around this tube, the complete satellite bus was developed to be "a wrap-around design of total volume 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 using technologies from the consumer electronics and automotive sector. For attitude determination, the satellites feature a star camera, global positioning system (GPS), photo-diodes for coarse sun sensing, and a micro-electro-mechanical system (MEMS) inertial measurement unit (IMU). Attitude control is 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-S-band downlink transmitter. A low-power processor equipped with solid state storage is running an unix-like operating system. Boshuizen et al. in [6] state, that none of these components "would have fit in 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 and
independent universities to highlight the difference between a CubeSat bus developed completely by one university over a bus commercially-procured by an university.
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 from Clyde 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 of Danish supplier GomSpace [32]. The satellites of the Kepler Communications constellation are based on the Clyde Space triple unit CubeSat bus [33]. And Aerial & Maritime is procuring the first four satellites of its planned constellation from Danish CubeSat company GomSpace. The satellite bus is based on the advanced GomSpace platform with heritage from the GOMX-3 mission, which is presented by Gerhardt, Bisgaard, and Alminde in [34]. But not only NewSpace companies procure CubeSat buses on the market. The spacecraft of NASA’s RAVAN mission as presented by Swartz et al. in [35] is based on Blue Canyon Technologies’s XB3 integrated triple unit CubeSat platform [36]. RAVAN is a technology demonstration mission for radiometry instruments used for analyzing the earth energy imbalance, which is a measure for 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 like Pumpkin, ISIS, SSTL, Clyde Space, or GomSpace are offering integrated plat-form solutions. To get a better understanding of those platplat-forms’ 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, with only three platforms aiming for single or double unit application. Besides the stated target CubeSat size, integrated platforms would be capable of supporting larger or smaller CubeSat form factors, with implications foremost on attitude control, due to the differing moment of inertia, and on payload power, due to solar cell area and therefore available power.
2.1.2.1 Payload Volume
Single unit platforms allow for a maximum payload volume ratio of 25 %. In comparison 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 is explained by the fact, that single and triple unit platforms of the same vendor rely on a similar set of CubeSat subsystems. For compatibility reasons with other suppliers, those subsystems are implemented on individual printed circuit boards (PCBs) and conform to the PC/104 form factor, which is the de-facto standard 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 kg mass 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 of CubeSat deployers, however, have establisehd a 2 kg per unit maximum mass limitation. 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 Source
Volume Volume Mass Power
U U kg W Clyde 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]
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, single and double unit CubeSats achieve a rather homogeneous payload mass ratio of 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 aluminum housings used for all subsystem to increase radiation tolerance [40], making them significantly heavier than the subsystems of most other companies. Triple unit payload mass ratios show a wide spread, spanning from 33.3 % for the Cube-X [44] to 75 % for the XB3 platform. Due to poor documentation in 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 with body-mounted solar cells the available specific power increases for smaller satellite masses. They explain this with the fact, that satellite mass is related to the outer satellite dimensions with the third power, while available solar cell area is related only with the second power to the outer dimensions. For satellites with a mass between 1–2 kg, Bouwmeester and Guo document a maximum specific power of 2 W/kg. It drops to about 1 W/kg for satellites in 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, which agrees well with the specific powers presented by Bouwmeester and Guo in [47]. The situation for double and triple unit spacecraft appears to be very inhomogeneous, with available payload powers between less than 1 W/U and 21.7 W/U. Expressed as specific power, this would be in a range of up to about 16 W/kg.
Observed available payload power of single unit spacecraft is well explained by the 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
also well explained with the application of deployable solar panels. Smaller values of less than 4 W/U were observed for several platforms. A closer look at the original data sources shows, that except for the ISIS basic bus, the given average payload powers are the absolute minimum values, and may easily 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 and magnetometers are the most commonly used sensor type within all pico- and nanosatellites between 1957 and 2009. Nearly 30 % of all missions in their database were equipped with either one or both of these sensor types. Funke et al. in their 2016 analysis on the characteristics and development of small satellites [48] state, that coarse sun sensors, magnetic field sensors, and angular rate gyroscopes have become the state of the art for small satellite attitude determination. To achieve precision attitude knowledge, however, additional star cameras are necessary.
The column for attitude knowledge data gathered in table 2.2 shows the biggest gaps. Only three triple unit platform manufacturers explicitly state the attitude knowledge. Tyvak’s is ranked top among the others with few arcsec attitude knowledge, while Clyde Space’s and GomSpace’s platforms support 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 systems with 5° and above pointing accuracy, and high-precision systems that feature a pointing accuracy of better than 1°. A similar distinction is made for agility, represented by the slew rate of maneuvers. While GomSpace’s single unit CubeSat platform has only a slew rate of 0.17 °/s, all triple unit platforms feature 3 °/s and above. The larger slew rates of the triple unit platforms are achieved using sets of at least three reaction wheels.
With a pointing accuracy of 7.2 arcsec and a slew rate of 10 °/s, Blue Canyon Technologies’s integrated triple unit CubeSat platform stands out from all other platforms. This performance is achieved using reaction wheels and a sensor 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 host UHF communications equipment, and achieve up to 19.2 kbps data rate. The majority of advanced configurations feature at least S-band communication, which starts at a data rate of 3.40 Mbps. The high-end variants offer additional X-band communication, with data rates up to 800 Mbps in the case of the Clyde 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 most commercial developers provide their subsystems with this interface". Their survey showed, that a growing number of CubeSats to be launched will implement the PC/104 connector, which they justify with the increasing availability of commercial subsystems that feature the 104-pin stackable connector. They further conclude, that a small majority of participants in their survey states that the PC/104 connector is too big, while otherwise no major problems are seen with the standard.
Bouwmeester, Langer, and Gill in [38] recommend that a future standard interface connector for CubeSat subsystems should be smaller than the PC/104 connector. They further conclude that this standard interface should have a fixed pin allocation to achieve general compatibility between subsystems.
Table 2.2:Integrated CubeSat platforms: attitude determination and control system characteristics
Brand Satellite ADCS Source
Volume Knowledge Accuracy Slew
U arcmin ° °/s Clyde Space 1 5 [33] GomSpace 1 0.17 [32, 39] Space Inventor 2 BCT 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 3 Tyvak 3 0.043 0.057 3 [45]
Table 2.3: Integrated CubeSat platforms: communication system characteristics
Brand Satellite Communication Source
Volume Band Rate
U kbps
Clyde 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 3
2.1.3 Independent University Spacecraft
Flagship and prolific independent universities are the groups that dominate the educational CubeSat launches, as stated by Swartwout and Jayne in [22]. The first flagship university listed in their list of spacefaring universities happens to be TU Berlin. Following the 2016 launch of BEESAT-4 [49], the 2017 launch of TechnoSat [50], and the 2018 launch of the S-Net formation [23], TU Berlin has launched a total of 16 small satellites over the last 27 years. With the upcoming launches of the BEESAT-5 to BEESAT-8 [26] formation and TUBIN [51] over the next two years, the number of small satellites launched by 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 spacecraft BEESAT-4. Its design is based on the BEESAT and BEESAT-2 missions and the satellite was deployed from the larger, German Aerospace Center (DLR)-operated, BIROS spacecraft after four months in orbit. Primary mission objective is the in-orbit demonstration of the Phoenix nanosatellite GPS receiver [52]. Most of the information on the BEESAT-4 satellite in this section 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 and angular rate sensors for attitude determination. Three RW 1 reaction wheels are used for three-axis attitude stabilization. Six magnetic coils support wheel angular momentum desaturation and spin dampening using a ˙𝐵 control law. The only subsystem not developed at or in cooperation with TU Berlin is the BK77 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 bus development was to achieve a complete as possible single-failure tolerance of every subsystem and therefore the complete satellite [54].
Electrical connection between the subsystems is not based on the PC/104 standard, a custom COTS connector solution is used. Communication between subsystems is happening on two independent controller area network (CAN) buses. The solar panels, which feature magnetorquers and TU Berlin’s precision sun sensors, are connected to the power control unit (PCU) board with insulation-displacement contact (IDC) connectors and flat ribbon cables [54]. All subsystems except for the PDH board, have flight heritage from BEESAT and BEESAT-2. The interior of the satellite is divided in two by the aluminum battery compartment. OBC, PCDU, and PDH boards occupy one half of the satellite. In the other half, the reaction wheels, wheel drive electronic (WDE) board, and two BK77 UHF transceivers are located. The two deployable antennas are also attached on this side of the spacecraft. Six solar panels complete the satellite, each hosting multiple solar cells, a sun sensor, and a magnetorquer. The panel close to the PDH hosts the GPS antenna. This one and 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 a camera 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 ratio of the 1 kg satellite is 7.18 %. Maximum power consumption of the PDH is 1.4 W [55]. However, not all payloads are operated simultaneously, and nominal 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 in comparison to other single unit platforms (cf. table 2.1), which allow for a ratio of 20–25 %. Owing to its heritage from BEESAT and BEESAT-2, the small payload volume ratio takes no wonder. For BEESAT, where the payload consisted of the reaction wheel assembly (RWA), WDE, and the PDH board including the camera, the estimated payload volume ratio is about 35–40 %.
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 mass ratio 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 𝑙
2, (2.1)
where 𝑀 is satellite mass and 𝑙 is the edge length of the satellite. Doubling the mass from 1 kg to 2 kg would therefore lead to double the inertia, cutting the maximum achievable angular rate in half. On-orbit results for attitude control data presented by Weiß and Kapitola in [53] document, that the spacecraft performs slew maneuvers at an angular rate of 5 °/s. During this maneuver, the reaction wheels reach a speed of about 2 500 rpm, which is only about 31.3 % of the nominal, and 15.6 % of the maximum wheel speed. This underlines, that the existing BEESAT satellite bus would be able to host heavier payloads. Weiß and Kapitola in [53] state, that the active attitude control operational time is currently limited by the power budget to a maximum of 60 min. Therefore the power budget would be the limiting factor to increasing the overall mass of the existing satellite.
Comparison of available payload power is more difficult to conduct. Maximum payload power consumption stated on page 18 is not representative for average power consumption at nominal satellite operation with GPS and inter-satellite link (ISL) experiments (cf. [53]).
With an attitude error of about 5° based on sun sensor and magnetic field measurements, BEESAT-4 is comparable to basic CubeSat platforms (cf. table 2.2). Slew rates of 5 °/s are positioning BEESAT-4’s agility among the higher 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 at the lower end of the integrated platforms (cf. table 2.3). Therefore, the communication system is a bottleneck for possible high performance payloads on the existing BEESAT bus.
The BEESAT satellites do not rely on subsystems that conform to the PC/104 standard. This offers flexibility in terms of harness and connectors, and allows
to put the reaction wheels on a common baseplate in one sixth of the volume of the satellite, while the WDE and two UHF transceivers are occupying the adjacent sixth of the internal volume. An assembly using subsystems conforming to PC/104, like e.g. the attitude determination and control system (ADCS) of GOMX-3 which uses a tetrahedron RWA based on four RW 1 reaction wheels and the associated WDE board mounted on a larger motherboard [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 the Universität 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 the re-utilization of a system, which is not inherently designed to be extended, can hardly be upgraded without producing a significant increase in total system complexity". For the design philosophy of UWE-3 and his successor missions, this lead to a preservation of the authority over the subsystem interface, thus relying on third party products only on the component level. The following discussion 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. All subsystem boards are directly connected to the backplane using standardized connectors. This connector implements the subsystem interface, which com-bines power lines for different voltages, GPIO, and digital interfaces like UART or I2C. In addition, dedicated lines for global reset and debug support are located on the electrical bus. Malfunctioning subsystems may be completely powered down to protect the digital buses from undesired bus allocation. Umbilical connections and the remove before flight (RBF) pin are realized using the first board connected to the backplane, the so-called front access board (FAB), which holds the required external connectors. All solar panels are attached directly to either the backplane or the FAB [57].
While the backplane and the FAB have advantages in terms of defined interfaces as well as easy access and testability, they are very sparsely populated with
electric components and thus occupy volume, that could be used by other subsystem boards or payloads. Also the support of multiple digital interfaces like universal asynchronous receiver transmitter (UART) or inter-integrated circuit (I2C), of which some were not designed to be used in harsh environments and therefore drive the need for additional components for selective bus isolation, is criticized in comparison to the use of a single, robust interface that was designed for harsh environments, like e.g. CAN bus. The use of a backplane has the further disadvantage that late changes to the arrangement of subsystems are rendered impossible owing to the need of having a new backplane PCB manufactured and equipped with connectors.
UWE-3 uses spacers in the board stack together with four identical rails as base for the mechanical structure. The solar panels, consisting of a two-layer PCB 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 in an integrated compact and lightweight design" [56].
Due to the two-layer stack-up and the aluminum core, it is difficult to increase the component count on the UWE-3 solar panels. In particular as the air-cored coils further limit the available surface area on the panels.
The UWE-3 bus, in conclusion, offers major advantages over existing CubeSat buses in terms of ease of assembly, subsystem interfacing, and robustness. Yet, there is seen potential for further advancing the bus concept and miniaturizing the volume required for the backplane and FAB.
2.1.3.3 Miniaturized Components for Distributed Pico- and Nanosatellite Systems
In the scope of the PiNaSys and PiNaSys II projects at TU Berlin, researchers aim at further miniaturizing CubeSat components to enable distributed satellite systems. Baumann et al. are developing a swarm consisting of four quarter-unit CubeSats, named BEESAT-5 to BEESAT-8, that will be deployed together from a single unit CubeSat deployment container in 2018 [28].
To miniaturize the satellite to one quarter of an unit, all satellite subsystems are integrated on one PCB, except for the batteries, which are mounted inside the metallic structure of the satellites close to the center of mass. Almost complete single point of failure tolerance is achieved by using the same board twice, and connecting them via a redundant CAN bus. The primary mission concept is to demonstrate the miniaturized bus, which is equipped with a number of novel, highly miniaturized components for communication and attitude 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 the panel. Internal communication between those subsystems is organized via a redundant CAN bus that is managed by the PCU [26]. The other side of the PCBs 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 increases complexity during development, as multiple engineers need to work on the same PCB design to integrate their subsystems. Development cost for suc-cessor missions with different mission objectives is increased due to the large amount of work for PCB design changes.
As the complete satellite is an in-orbit demonstration (IOD) mission, it is difficult to define a ratio between payload and bus volume and compare this to commercially procured CubeSats based on integrated platforms (cf. section 2.1.2). From the subsystem point of view, the BEESAT-5 to BEESAT-8 satellites will demonstrate high performance attitude and orbit determination capabilities using the experimental star cameras and GPS receivers. As magnetorquers are the only means for attitude control, precise and agile pointing is out of range. The X-band transmitter provides payload data downlink 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 of electric energy to achieve the high data rates.
2.2 High Performance Single Unit CubeSat Design
Criteria
From the discussion of the characteristics of commercially-procured and independent university spacecraft in sections 2.1.2 and 2.1.3, respectively, a certain set of design criteria for a high performance single unit CubeSat are derived. Compatibility with existing solutions offered on the CubeSat market is not a requirement.
2.2.1 Provided Payload Resources
The ratio between available payload volume and overall CubeSat volume does not scale for single unit CubeSats in comparison to the larger CubeSats: While triple unit spacecraft have at least 50 % of volume available for payloads, single unit CubeSats allow only for up to 25 % (cf. section 2.1.2.1). Main reason for this is the utilization of CubeSat subsystems that obey to the PC/104 standard (cf. [38]). Avoiding PC/104-based subsystems and exploiting the level of miniaturization seen in the scope of the PiNaSys and PiNaSys II projects, a better volumetric utilization of about 50 % with available payload masses as high as 1 kg should be achievable.
Section 2.1.2.3 documents 2 W average payload power available on single unit platforms. A general increase in bus performance leads to an increased bus power demand. Keeping 2 W power available for the payload therefore might require 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. For pointing accuracy, a distinction is made between low-precision platforms with 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
MEMS magnetic field and angular rate sensors for attitude determination and reaction wheels with magnetorquers for attitude control. Reaction wheels allow for larger slew angles and the availability of control torque about all three 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 platforms reveals, that they currently rely on UHF communication with data rates up to 19.2 kbps (cf. table 2.3). The majority of triple unit spacecraft feature S-band with data rates starting from 3 400 kbps. However, no data on power demand of the featured communication hardware is provided. For a power-limited single 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 designed for CubeSat applications. Utilization of a S-band transmitter drives the need for 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 the PC/104 standard. A backplane featuring identical interface connectors for every 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 the payload-to-bus ratios. While using a backplane, the modular TUBiX20 plat-form developed at TU Berlin offers a common data and power bus where all inter-subsystem communication is carried out on a redundant CAN bus interface (cf. [60]). This approach to modular satellite architecture allows to adapt the platform to new missions with a minimum effort in terms of redesign of existing hardware, while it allows for full flexibility in adding or removing functionality.
The CubeSats of TU Berlin, like BEESAT-4 and the upcoming BEESAT-5 to BEESAT-8 satellites, already feature several subsystems, that are connected
