6th European CubeSat Symposium
14 – 16 October 2014
Estavayer-le-Lac, Switzerland
Book of Abstracts
Sponsors:
2
Table of Contents
Foreword…………………………………………………………………………………….....
3
Organisational details……………………………………………………………….........
4
Programme........................................................................................
7
Scientific instruments/sensors on CubeSats...........................................
14
Micropropulsion Systems……………………................................................ 30
CubeSat Launchers and Deployers……………………………….......................
35
Technology Demonstration on CubeSats...............................................
39
Telecommunications, Ground Stations, Ground Station Networks………….
47
CubeSat Flight Experience, Lessons Learned…………..............................
58
Orbital Dynamics, De-orbiting and Debris Mitigation Techniques.............
67
CubeSat Networks and Constellations, Formation Flying......................... 73
Attitude Determination and Control ……………........................................ 77
Future Technologies on CubeSats.......................................................
87
Posters...............................................................................................
94
3
Foreword
When the European CubeSat Symposium was first organised, CubeSats were seen
as dominantly educational tools that served for the training of university students.
Today, while we are organising the 6th European CubeSat Symposium, we are also
congratulating hundreds of CubeSats launched to orbit. Only in 2014, two
spectacular launch campaigns have been performed: First, 33 CubeSats were
launched in January 2014, and then 24 others in June 2014. A significant number of
these CubeSats were developed by Planet Labs for commercial Earth Observation
purposes, showing that CubeSats are now used for commercial, scientific and
technology demonstration reasons, in addition to being educational tools. Launching
of CubeSats from the International Space Station has been another recent
pioneering step. In addition to the classical 1, 2U and 3U concepts, we also started
seeing designs of bigger nano and micro satellites based on 6U and 12U CubeSat
concepts.
The CubeSat community is growing fast parallel to all these developments. The 6th
European CubeSat Symposium has attracted full attention from the community with
more than 100 abstracts submitted from 31 different countries. Von Karman Institute
and Swiss Space Systems are proud to support the CubeSat community by coorganising this leading CubeSat event in Europe, for the first time in Switzerland.
Von Karman Institute continues to act as the coordinator of the World's most
ambitious CubeSat Project QB50, whereas Swiss Space Systems is designing an
innovative launcher specifically for small satellites to bring the launch costs to 25% of
today's market value.
We are very happy to have contributed to this year's CubeSat Symposium where 70
oral presentations will be given in 11 different sessions. 8 industrial companies will
be presenting their products and solutions in a parallel industry session and at the
industrial exhibit. 19 posters will take place in the poster sessions, waiting for
detailed discussions during coffee breaks, lunch and reception. The event is fully
booked and we are expecting more than 200 participants. The 6th European
CubeSat Symposium has been a great success so far by attracting several sponsors
such as the European Space Agency, Tyvak Nano Satellite Systems Inc, Journal of
Small Satellites, Swiss Space Systems, ClydeSpace and Aeropole/COREB to
sponsor more than 12 participants. Another interesting number is the 9 presentations
in the “CubeSat Flight Experience, Lessons Learned” session, which is attracting
more and more presentations every year.
We wish you a very successful Symposium and look forward to meeting you again
next year at the 7th European CubeSat Symposium.
Pascal Jaussi
Jean Muylaert
CEO
Director
Swiss Space Systems
von Karman Institute
4
Organisational details
Access to the premises
The Symposium takes place at the Théâtre de la Prillaz, in the city of Estavayer-le-Lac in Switzerland.
The open address is:
Chaussée des Autrichiens 15
CH-1470 Estavayer-le-Lac
Switzerland
Estavayer-le-Lac is a town on the eastern coast of the Lake of Neuchatel. It has direct train
connections to bigger cities such as Yverdon and Fribourg. The closest airport is the Airport of
Geneva. The train connections can be checked and tickets can be purchased at
http://www.sbb.ch/en/
There will be shuttle services between the Symposium venue and the train station of Estavayer-leLac, as well as between the Symposium venue and Hotel Park Inn Lully. There is a big parking place
at the Symposium venue for those who travel by car.
5
Registration
The registration desk will be open at the entrance of the Symposium venue from 08.45 to 09.45.
However, late arriving participants can still register. Participants will be asked to pay a modest
registration fee of 250 € to cover the expenses for lunches and coffee breaks on three days and for
drinks and snacks during the reception.
Oral presentations, proceedings
Speakers will not be allowed to use their own computer during their talk, but must transfer their
presentation in PDF or PPT PowerPoint to the Symposium Secretary (cem.asma@s-3.ch) by email or
by USB flash drive, preferably half a day before the presentation.
There will be no printed Symposium proceedings, only the abstracts will be published. The slides of
all presentations will be made available after 20 October 2014 on the Symposium website.
Posters
The size of the area available for poster presentation is 120 x 85 cm. A1 size posters can be exhibited
during the Symposium. The accepted participants can bring their own posters or the posters can be
printed by Swiss Space Systems in advance (this should be the exception). Poster mounting will be
possible on 14 October 2014 from 08.30 until 11.30. Standard materials for poster mounting will be
available in the poster area, where the coffee breaks and receptions will be held. Poster authors are
expected to be present at that time next to their posters, to be available to answer questions and
have discussions with Symposium participants.
Industrial exhibits
There will be industrial exhibits by the following companies:
 Blue Canyon Technologies, USA
 Bright Ascension, UK
 Clyde Space, UK
 GomSpace, Denmark
 GOSMOZ, Switzerland
 Innovative Solutions in Space, Netherlands
 Tyvak Nano-Satellite Systems, USA
 Observatoire de Versailles Saint-Quentin-en-Yvelines (OVSQ), France
The industrial exhibits will take place in the poster/coffee/reception area.
Lunch breaks
The lunch break is between 12.30 and 14.00 every day.
Coffee breaks
The times for morning and afternoon coffee breaks are indicated in the programme. The participants
will also have the possibility to purchase beverages and drinks outside the coffee breaks.
Reception
On the first day of the Symposium, from 19.00 to 20.30, a reception will take place in the
poster/coffee/reception area (kindly sponsored by Aeropole.ch).
A goodbye drink will be held at the last day of the Symposium (kindly sponsored by ClydeSpace).
6
International Scientific Committee
W. Balogh (UN)
R. Walker (ESA)
O. Koudelka (Austria)
J. Muylaert (Belgium)
J. Thoemel (Belgium)
J. F. Dalsgaard Nielsen (Denmark)
K. Briess (Germany)
R. Reinhard (Germany)
E. Gill (Netherlands)
J. Rotteveel (Netherlands)
T. Masson-Zwaan (Netherlands)
S.R. Cunha (Portugal)
V. Gass (Switzerland)
P. Jaussi (Switzerland)
V.I. Mayorova (Russia)
D. Kataria (United Kingdom)
G. Shirville (United Kingdom)
T. Morgensen (USA)
S. Palo (USA)
Symposium Secretaries:
C. O. Asma, Swiss Space Systems
D. Masutti, von Karman Institute
P. Testani, von Karman Institute
7
Programme
Tuesday, 14 October 2014
08.45 – 09.45
Registration, Coffee
09.45 – 10.00 Welcome (C. Leu and J. Thoemel) and Opening Speech: Micro, Nano and Small
Satellites in the context of the Swiss Space Policy (J. Richard – Swiss Space Office)
10.00 – 10.15
Practical Information (C.O. Asma)
Scientific Instruments/Sensors on CubeSats
Chair: D. Masutti
10.15 – 10.30 Interfacing with the science unit: Preparing the software side
(R. A. de Carvalho et al.)
10.30 – 10.45 The FIRST-S Project : the science case (S. Lacour et al.)
10.45 – 11.00
Antenna development for the Wideband Ionospheric Sounder CubeSat Experiment
(WISCER) (G. C. Kirkby and M. J. Angling)
11.00 – 11.15
A compact ion and neutral mass spectrometer for the ExoCube Mission
(N. Paschalidis et al.)
11.15 – 11.30 The PIC.A.S.S.O. mission: A PICo-satellite for Atmospheric and Space Science
Observations (D. Fussen et al.)
11.30 – 11.45
Solar EUV Probe on QB50/PHOENIX CubeSat (J. C. Juang et al.)
11.45 – 12.00
CubeSat-ready radiation monitor front-end electronics (T. A. Stein)
12.00 – 12.15
Wavelet transform as an efficient and effective tool for electromagnetic emission
measurements in space plasma (T. Szewczyk et al.)
12.15 – 12.30 On feasibility of Global Radio Frequency Interference measurement with CubeSats
(J. Praks et al.)
__________________________________________________________________________________
12.30 – 14.00 Lunch break
__________________________________________________________________________________
8
Scientific Instruments/Sensors on CubeSats (continued)
Chair: A. Ridley
14.00 – 14.15 Doing Science with university Cubesats: Present and future (T. Moretto Jorgensen)
14.15 – 14.30 The micro solar-flare apparatus (MiSolFA) (D. Casadei)
14.30 – 14.45 Mapping the radiation environment in equatorial low earth orbit with CubeSats
(E. Del Monte et al.)
14.45 – 15.00 From the low cost to the high benefits. The outreach of small satellites
instrumentation (M.D. Michelena )
15.00 – 15.15 A CubeSat for X-ray polarimetry in astrophysics (P. Soffitta et al.)
15.15 – 15.30 Evaluation of the performances of a spectro imaging detector (CALISTE) to be
embedded on board a nanosatellite for Solar Flares studies (H. Triou et al.)
__________________________________________________________________________________
15.30 – 16.00 Coffee break
__________________________________________________________________________________
Micropropulsion Subsystems
Chair: M. Richard
16.00 – 16.15 A highly miniaturized uPPT thruster for attitude-orbit control (J. Li et al.)
16.15 – 16.30 Solid cool gas micro propulsion system for CubeSat (P. Zhu et al.)
16.30 – 16.45 Solid propellant micro-thruster array for CubeSat (X. Liu et al.)
16.45 – 17.00 Development status of an open capillary pulsed plasma thruster with non-volatile
liquid propellant (E. Remírez et al.)
CubeSat Launchers and Deployers
Chair: M. Richard
17.00 – 17.15 Reliable low-cost accesses to space for CubeSat size payloads (P. C. Steimle et al.)
17.15 – 17.30 On-orbit optimization of CubeSat launch opportunities using an orbital maneuvering
vehicle (M. Stender et al.)
17.30 – 17.45 Swiss Space Systems innovative launch concept for small satellites
(B. Deper et al.)
9
18.00 – 19.00 Round Table Discussions: Innovation in Space
Beat Vonlanthen, President, State of Fribourg
Pascal Jaussi, CEO & Founder, Swiss Space Systems
Clément Leu, Director of Education & Outreach, Swiss Space Systems
__________________________________________________________________________________
19.00 – 20.30 WELCOME RECEPTION
(kindly sponsored by Aeropole.ch)
__________________________________________________________________________________
10
Wednesday, 15 October 2014
Technology Demonstration on CubeSats
Chair: A. Denis
08.45 – 09.00 Space object detection on a CubeSat platform (M. Tetlow et al.)
09.00 – 09.15 The Multi-Payload Satellite Service (MPS) for in-orbit technology evaluation
(L. León et al.)
09.15 – 09.30 SamSat–QB50 nanosatellite: Burning wires mechanism for antenna system and
aerodynamic stabilizer (E. Ustiugov et al.)
09.30 – 09.45 Testing of low-cost GNSS receivers for CubeSat orbit and attitude determination
(C. Hollenstein et al.)
09.45 – 10.00 QARMAN: As an Atmospheric Entry Experiment on CubeSat Platform (E. Umit et al.)
10.00 – 10.15 A thermal protection system for a re-entry CubeSat (P. Testani et al.)
10.15 – 10.30 A tool for nano-satellite functional verification: comparison between different inthe-loop simulation configurations (L. Feruglio et al.)
__________________________________________________________________________________
10.30 – 11.00 Coffee break
__________________________________________________________________________________
11.00 – 11.30 Invited Presentation: CubeSat Missions at Morehead State University - From LEO to
Lunar (Dr. Benjamin K. Malphrus, Director of the Space Science Center, Morehead
State University)
Telecommunications, Ground Stations, Ground Station Networks
Chair: A. Denis
11.30 – 11.45 Development of an open source software defined radio (M. Wegerson et al.)
11.45 – 12.00 CubETH COM subsystem (F. Belloni et al.)
12.00 – 12.15 Enhancements of antenna control and an error correction scheme for the CubETH
Ground Station (S. Kaufmann et al.)
12.15 – 12.30 SUSat communications system (W.G. Cowley et al.)
__________________________________________________________________________________
12.30 – 14.00 Lunch break
__________________________________________________________________________________
11
Telecommunications, Ground Stations, Ground Station Networks (continued)
Chair: M. Joss
14.00 – 14.15 Antenna subsytem of GAMALINK platform (V. Akan and C. Dudak)
14.15 – 14.30 A simple miniaturized printed antenna adaptation for CubeSats and small satellites
(S. Kose et al.)
14.30 – 14.45 Development of an x-band transmitter for CubeSats (S. Palo et al.)
14.45 – 15.00 Multi-objective optimization of a high gain, circularly polarized rectangular antenna
array in the Ka band for CubeSat class satellites (A. Cuttin et al.)
15.00 – 15.15 KSAT light – a low cost ground station network for Cubesats
(B. Eilertsen and M. Krynitz)
15.15 – 15.30 Challenges and solutions for the QB50 telecommunication network (G. March et al.)
__________________________________________________________________________________
15.30 – 16.00 Coffee break
__________________________________________________________________________________
CubeSat Flight Experience, Lessons Learned
Chair: J. Thoemel
16.00 – 16.15 Lessons learned from four months of UKube-1 in orbit: A software perspective
(P. Mendham and M. McCrum)
16.15 – 16.30 The QB50 precursor flight: Lessons learned (J. Elstak et al.)
16.30 – 16.45 QB50 precursor ADCS flight results (L. Visagie et al.)
16.45 – 17.00 A worldwide survey on the regulatory and economical aspects of nano-satellites
(S. Cabrera et al.)
17.00 – 17.15 LituanicaSAT-1: lessons learned from the first Lithuanian cubesat mission
(L. Maciulis et al.)
17.15 – 17.30 Lessons learned from developing and producing structure and mechanical systems
for ESTCube-1 (P. Liias et al.)
17.30 – 17.45 Legal Aspects on CubeSats and Space Debris Issues (N. Antoni)
17.45 – 18.00 E-st@r-I lessons learned and their application (G. Obiols-Rabasa et al.)
12
Thursday, 16 October 2014
Orbital Dynamics, De-Orbiting and Debris Mitigation Techniques
Chair: S. Palo
08.30 – 08.45
The Aerospace Blockset for Xcos (P. Zagórski)
08.45 – 09.00
LitSat-1 decay analysis (V.Tomkus et al.)
09.00 – 09.15
In orbit testing of a de-orbiting sail on the Cubesat URSA MAIOR (M. Valdatta et al.)
09.15 – 09.30
A tether-based aerodynamic de-orbiting system (O. Vallet and C.O. Asma)
09.30 – 09.45 De-risking active debris removal with CubeSat in-orbit demonstrations
(M. Richard et al.)
CubeSat Networks and Constellations, Formation Flying
Chair: S. Palo
09.45 – 10.00 Space-based ad hoc network: a solution for multiple satellite TT&C problem in QB50
project (P .Liu et al.)
10.00 – 10.15
TW-1: A Cubesat constellation for space networking experiments (S. Wu et al.)
10.15 – 10.30
Status of the QB50 Project (J. Thoemel et al.)
__________________________________________________________________________________
10.30 – 11.00 Coffee break
__________________________________________________________________________________
11.00 – 11.30 Invited Presentation: Market Evolution and Commercialization of CubeSats (Dr.
Marco Villa, President and COO of Tyvak Nano-Satellite Systems, Inc.)
Attitude Determination and Control Subsystem
Chair: J. Praks
11.30 – 11.45 Aalto-1 nanosatellite attitude determination and control system end-to-end testing
(T. Tikka et al.)
11.45 – 12.00 Star tracker cost reduction for small satellites (T. Delabie et al.)
12.00 – 12.15 ZA-AeroSat: A QB50 CubeSat demonstrator for multidisciplinary technology and
scientific research (M. Kearney and W.H. Steyn)
12.15 – 12.30 Attitude control simulation using variable speed CMG for 3U CubeSat (H. Kim et al.)
__________________________________________________________________________________
12.30 – 14.00 Lunch break
__________________________________________________________________________________
13
Attitude Determination and Control Subsystem (continued)
Chair: R. Atem de Carvalho
14.00 – 14.15 The piNAV-L1 – The World’s first ultra low power CubeSat GNSS receiver (J. Laifr)
14.15 – 14.30 In-house magnetic field simulator for Cubesats (M. Balan et al.)
14.30 – 14.45 Active magnetic attitude control algorithms for CXBN-2 CubeSat
(M. Ovchinnikov et al.)
14.45 – 15.00 Characterisation of hysteretic dampers for passive attitude control of Cubesats
(D. Ivanov et al.)
15.00 – 15.15
A constrained attitude control method for Aoxiang-Sat (R.Liu et al.)
__________________________________________________________________________________
15.30 – 16.00 Coffee break
__________________________________________________________________________________
Future Technologies on CubeSats
Chair: M. Tetlow
16.00 – 16.15 Characterization and design of solid state hinges for deployable Cubesat structures
(E. Ziade et al.)
16.15 – 16.30 Autonomous command and data handling system for a 3U CubeSat
(L. Feruglio et al.)
16.30 – 16.45 Early orbit phase of deployment mission of inflatable membrane structure of nanosatellite ''SPROUT'' (K. Mita et al.)
16.45 – 17.00 Implementation of an on-board computer & a modem into a single subsystem for
CubeSat (M.E. Bas et al.)
17.00 – 17.15 BIRDY: An interplanetary CubeSat to collect radiation data on the way to Mars with a
precursor flight around the Earth in GTO (B. Segret et al.)
17.15 – 17.30
Target Shape Identification for Nanosatellites using Monocular Point Cloud
Techniques (Mark Post and Xiu-Tian Yan)
17.30 – 17.45 Closing remarks (P. Jaussi & J. Muylaert)
__________________________________________________________________________________
17.45 – 19.00 GOODBYE DRINK (kindly offered by Clyde Space)
__________________________________________________________________________________
14
Scientific Instruments/Sensors
on CubeSats
15
Interfacing with the Science Unit:
Preparing the Software Side
R. A. de Carvalho1, H. S. Ferreira1, R. F. Toledo1, C. S. Cordeiro1 and Moura L.G.L.1
1
Instituto Federal Fluminense, Campos dos Goytacazes (Rio de Janeiro), Brazil
All QB50's satellites will carry and operate one of the three different sensors for in-situ
measurements defined by the mission coordination. The development of these sensors is in parallel
with the development of the satellites that will transport them, through “contracts” established in
proper Interface Control Documents (ICDs), in such a way that satellite developers can prepare their
spacecraft to receive and operate the sensors. The aim of this paper is to present the environment
created by 14-BISat Team to test the software that will control the Fipex (Flux-Φ-Probe Experiment)
through its scripts, store the collected data, and forward this data through GAMALINK, the S-Band
Software Defined Radio onboard of 14-BISat.
In order to provide a realistic simulation of the Fipex functioning, while aiming at the quality of the
code, a set of elements was used to form an integrated software development environment: (i) a
tool for modelling and simulating Finite State Machines (FSM), (ii) an Integrated Development
Environment for micro-controllers, (iii) a C programming language Test Harness, and (iv) the
simulation of Fipex interfaces using a MSP430 Launchpad. The proposed environment provides an
integrated process for testing the software on top of the target sensor interface definitions: draw
the FSM that represents the sensor behaviour; create test cases for regular and problematic
behaviour; automatically generate the tests skeletons for the test cases - with their expected results;
and run and check the results. Aiming at supporting this test process, it was developed: a plug-in to
generate the automated tests skeletons, hooks for configuring all the tools to work in an integrated
way, a library to simulate I/O problems, and a simulation of the Fipex command and response
mechanisms for the micro-controller.
The proposed process and its supporting environment are capable of testing the software up to the
limits that the absence of the (real) hardware allows. Moreover, this set of practices and tools can be
generalized to the development of other embedded softwares.
16
The FIRST-S Project: the science case
S. Lacour1, V. Lapeyrere1, L. Gauchet1, S. Arroud1, G. Ronnan1, and G. Perrin1
1
Laboratoire d'Etudes Spatiales et d'Instrumentation en Astrophysique (LESIA),
Observatoire de Paris, France
The FIRST-S project is an astronomical project in the context of exoplanet detection. The goal is to
measure the amount of exozodiacal light scattered by dust around bright nearby stars. The level of
zodiacal light is a problem since it can hamper planetary detection in extrasolar systems.
In the near infrared, the thermal emission of the exozodiacal dust has already been observed from the
ground. It is typically 1% of the stellar flux. But the question of the level of scattered light by the dust in
the habitable zone is still a mystery. At 1 AU of the star, it may be several thousands times fainter. Still,
integrated over the exoplanetary system, it may be several thousand times brighter than the flux
emitted by the planets.
This is why it needs a dedicated project. The problem from the ground is not the angular resolution. A 8meter telescope theoretically reach an angular resolution equivalent to 0.12 AU at visible wavelength
(for a star at 10pc). The problem is the dynamic range, which is limited by the atmosphere, even with an
active optics system.
Our goal is therefore to reach high dynamic range from space, at moderate resolution. To do so, the
FIRST-S CubeSat will be a 30 cm stellar interferometer, with broadband nulling capabilities of the order
of 10^3. It is currently under development.
17
Antenna development for the Wideband Ionospheric Sounder CubeSat
Experiment (WISCER)
G. C. Kirkby, M. J. Angling
Space Environment and Radio Engineering group, University of Birmingham, UK
There are a wide range of potential uses for a space based foliage penetrating (FOPEN) synthetic
aperture radar (SAR) system; however there are significant obstacles which must be dealt with in order
to develop an operational system. One of these is the impact of the ionosphere on the system.
Consequently, the Wideband Ionospheric Sounder CubeSat Experiment (WISCER) is being developed to
measure the impact of the ionosphere on a wideband radar-like signal. .
WISCER comprises a wideband (approximately 100 MHz) beacon operating at a centre frequency of
approximately 475 MHz. In combination with a ground station the system will measure the channel
impulse response (CIR) for the post sunset equatorial ionosphere. The post sunset equatorial ionosphere
is of particular interest due to the existence of small scale ionospheric structures that can cause
scintillation of the trans-ionospheric signal. It is this scintillation that results in degradation of SAR
performance (i.e. loss of contrast in SAR images).
It is necessary for the WISCER antenna to be efficient and to have some gain in the forward direction
due to the constrained power of the CubeSat. Furthermore, the antenna gain pattern should be well
behaved within the main antenna lobe and across the required bandwidth. This is necessary to ensure
that the ionospheric effects on the received signal are not obscured by the antenna characteristics.
These requirements lead to relatively large antenna designs which present difficulties for deployment
and for the CubeSat attitude control system (ACS).
Two antenna designs have progressed through initial design stages. The first is an inflatable conical helix
antenna. The antenna arms are printed on to the outer layer of the rigidising balloon. The second is a
mechanically deployed crossed-Moxon antenna with a re-purposed de-orbit sail acting as a ground
plane. This paper will present an overview of the two antenna designs and their electrical performance.
Furthermore, results from orbital disturbance simulations that bound the required authority of the
CubeSat ACS will be presented.
18
A Compact Ion and Neutral Mass Spectrometer for the ExoCube Mission
N. Paschalidis1, S. Jones1, M. Rodriguez1, E. Sittler1, D. Chornay2, P. Uribe1, T. Cameron1, G. Nanan1, G.
Suarez1, J. Dumonthier1, J. Noto3, L. Waldrop4, C. Taylor5, D. Gardner6, S. Nosal6, E. Mierkiewicz7
1
NASA / Goddard Space Flight Center (GSFC) , Greenbelt MD 20971, USA
2
University of Maryland, College Park, MD, US
3
Scientific Solutions Inc., MA, USA
4
University of Illinois at Urbana-Champaign
5
California Polytechnic State University, San Luis Obispo
6
University of Wisconsin-Madison
7
Embry-Riddle Aeronautical University, Daytona Beach, FL
Demand is high for in situ measurements of atmospheric neutral and ion composition and density, not
only for studies of the dynamic ionosphere-theremosphere-mesosphere system but simply to define the
steady state background atmospheric conditions. The ExoCube mission is designed to acquire global
knowledge of in-situ densities of [H], [He], [O] and [H+], [He+], [O+] in the upper ionosphere and lower
exosphere in combination with incoherent scatter radar ground stations distributed in the north polar
region . The Heliophysic Division of GSFC has developed a compact Ion and Neutral Mass Spectrometer
(INMS) for in situ measurements of ions and neutrals H, He, N, O, N2, O2 with M/dM of approximately
10 at an incoming energy range of 0-50eV, indented for ExoCube and other missions. The INMS is based
on front end optics, post acceleration, gated time of flight, ESA and CEM or MCP detectors. The compact
sensor has a dual symmetric configuration with the ion and neutral sensor heads on opposite sides and
with full electronics in the middle. The neutral front end optics includes thermionic emission ionization
and ion blocking grids, and the ion front end optics includes spacecraft potential compensation grids.
The electronics include front end, fast gating, HVPS, ionizer, TOF binning and full bi directional C&DH
digital electronics. The data package includes 400 mass bins each for ions and neutrals and key
housekeeping data for instrument health and calibration. The data sampling can be commanded as fast
as 10 msec per frame (corresponding to ~80 m spatial separation) in burst mode, and has significant
onboard storage capability and data compression scheme. Experimental data from instrument testing
with both ions and neutrals will be presented. The instrument is successfully integrated in the CubeSat
and passed vibration, thermal and shock testing. The ExoCube mission is scheduled to fly in Nov 2014 in
a 445 x 670 km polar orbit with the INMS aperture oriented in the ram direction. This miniaturized
instrument (1.5U), weighing only 560 gr and requiring peak power of 1.6W, will provide the first in situ
measurement of exospheric hydrogen and will measure in situ atomic oxygen for the first time in
decades.
19
The PIC.A.S.S.O. mission: A PICo-satellite for Atmospheric and Space Science
Observations
D. Fussen, D. Pieroux, S. Ranvier, J. De Keyser and P. Cardoen
Belgian Institute for Space Aeronomy, Brussels, Belgium
PIC.A.S.S.O. is a joint project led by the Belgian Institute for Space Aeronomy (BISA) in collaboration with
the Royal Observatory of Belgium (ROB). A triple-unit CubeSat targeting the QB50 flight will be
developed to embark two scientific experiments dedicated to the study of the ozone distribution in the
stratosphere, the temperature profile up to the mesosphere and the electron density in the ionosphere.
PICASSO falls thus within the category of CubeSats dedicated to specific scientific missions or to
technologic demonstrations, the so-called in-orbit demonstration (IOD) CubeSats.
Our goal in participating in a CubeSat mission is primarily to carry out actual scientific experiments, not
to develop innovative technologies, to demonstrate new concepts or to educate students.
To achieve that goal, PICASSO will embark two experiments:
1. VISION, a visible and near-infrared hyper-spectral imager
2. SLP, a Sweeping Langmuir probe;
BISA, which is responsible for both instruments, will delegate the realisation of VISION to the VTT
Company (Finland) and will internally develop SLP.
For the sake of efficiency and risk reduction, it has also been decided to entrust a CubeSat Industrial
Partner (Clyde Space, UK) with the development and the tests of the PICASSO platform from readily
available commercial-of-the-shelf (COTS) components. This partner will also be in charge of integrating
the instruments and preparing the mission. Optionally, it could be asked to operate the mission too.
PIC. A.S.S.O. offers an opportunity to assess the relevance of the CubeSat technology for atmospheric
soundings and in-situ measurements at a reduced cost,
20
Solar EUV Probe on QB50/PHOENIX CubeSat
J. C. Juang1, A. Chen2, C. W. Chao2, J. Vannitsen3, and J. J. Miau3
1
Department of Electrical Engineering, National Cheng Kung University, Tainan, Taiwan
Institute of Space and Plasma Sciences, National Cheng Kung University, Tainan, Taiwan
3
Department of Aeronautics and Astronautics, National Cheng Kung University, Tainan, Taiwan
2
The scientific objective of the QB50 is to study the temporal and spatial variations of ions, neutral
particles, electrons, and atomic oxygen in the lower thermosphere with a string of CubeSats carrying
sensors for in-situ measurements. Currently, sensors that will be deployed include Ion and Neutral Mass
Spectrometer (INMS), Flux--Probe (FIPEX), and multi-needle Langmuir Probe (m-NLP). The most
important energy source for the thermosphere is known to be the EUV and X-ray solar radiation.
Indeed, solar EUV radiation plays an important role in the ionization process. Therefore, in measuring
the contents of ions, neutral particles, and electrons, it is equally important to assess solar EUV radiation
intensity. As a participant of the QB50, the PHOENIX CubeSat which is under developed at National
Cheng Kung University (NCKU), Taiwan, is designed to carry a solar EUV sensor to in-situ measure the
solar EUV radiation intensity with the aim of maximizing the science return of the mission for a better
characterization of the cause-effect relationship. The solar EUV probe measures the photoelectron
current emitted from the electrodes into ambient plasma. The electrode is biased to negative voltage
with respect to the satellite frame and the current is amplified and filtered before being converted into
digital formats. In the implementation, two electrodes with different metal coating are utilized and the
current difference between the two probes. The solar EUV instrument has been tested in ionospheric
conditions in the NCKU Space Plasma Chamber. An integration test campaign has also been conducted
to check the functionality and performance of the sensor.
21
CubeSat-ready Radiation Monitor Front-End Electronics
T. A. Stein1,2
1
Department of Physics, Norwegian University of Science and Technology (NTNU),
7491 Trondheim, Norway
2
Integrated Detector Electronics AS (IDEAS), Martin Linges Vei 25, 1364 Fornebu, Norway
The requirements for CubeSats heavily restrict the payload in terms of mass, volume, and power. In this
work we present radiation monitor front-end electronics that meet the requirements for CubeSats. The
electronics design uses an application specific integrated circuit (ASIC) with multiple channels of charge
sensitive amplifiers and charge discriminators for pulse height spectroscopy and counting. The ASIC
provides a low-mass and low-volume solution for the entire front-end electronics on a monolithic silicon
die. The ASIC and the embedded system are optimised for space operations such as lowest possible
power consumption. In order to meet the CubeSat specifications, we adapted the voltage supply and
signal interface between the ASIC and on-board computer.
The ASIC-based design allows the spectroscopic counting of electrons and protons. The low mass and
small size makes the device ideal for CubeSats. Furthermore, the in-flight operation of the instrument in
CubeSats will be useful for further instrument developments that require flight heritage. Emphasis is put
into the generalisation of the device to meet the needs of the CubeSat community.
NTNU is currently developing a double CubeSat called NUTS aiming for launch in 2016. In view of future
CubeSat and small satellite missions the above front-end electronics were conceived within the NUTS
project in collaboration with IDEAS. The first flight of the CubeSat-ready Radiation Monitor Electronics is
proposed for the Norwegian satellite NORSAT-2. This mission is set for launch in 2018 and is currently in
its planning phase. The small satellite NORSAT-2 is going to carry several instruments predominantly
made by Norwegian institutions with opportunities for university students. We have conducted initial
tests of the ASIC and the results are very encouraging for its use towards a radiation monitor for the
NORSAT-2 mission.
We are analysing the implementation of the radiation monitor with the FPGA, the microcontroller and
the radiation detectors and the power system. These components allow relevant studies in fields such as
space weather or high-energy physics. We plan to use the device for the monitoring of electron and
proton radiation as well as energy spectroscopy.
22
Wavelet transform as an efficient and effective tool for electromagnetic
emission measurements in space plasma
Tomasz Szewczyk, Hanna Rothkaehl, Marek Morawski
Space Research Centre, Polish Academy of Sciences, Warsaw, Poland
Wavelet transform is successfully used in various areas of data analysis and compression (e.g. JPEG2000
image compression standard). Application of wavelets in analysis of the electromagnetic emission
observed by space borne instruments is currently investigated in Space Research Centre PAS. Wavelet
transform might be especially useful in:



Real-time filtration of data and detection of particularly interesting scientific activity.
Analysis both in time and frequency domain, and reconfiguration of instrument to focus on
certain scientific aspects of measured data.
Compression of gathered data allowing better utilization of communication bandwidth.
Implementation of new algorithms used in Earth orbiting frequency analyzers will allow better
understanding of on-going phenomena in ionosphere and magnetosphere. Process of hardware
implementation of wavelet transform toolset consists of the following steps:



Error analysis (in Matlab environment) of data/wavelet coefficient number representation. This
aspect is particularly important, since hardware implementation of wavelet transform has to be
robust, yet with small hardware footprint.
Verification of model using hardware VHDL implementation of wavelet transform in Xilinx Spartan3
FPGA. To provide reliable source of test data, signals gathered by Demeter satellite are used.
Hardware implementation of filtration and compression algorithms that were previously tested in
Matlab environment.
Wavelet processing toolkit is currently developed as one of scientific modes for High Frequency Analyzer
(HFA) instrument which will be placed on-board RESONANCE (constellation of four satellites) mission
whose aim is to perform 4-point measurements of Earth’s plasma. This mode is new feature in
electromagnetic wave analyzers developed at Space Research Centre. Further plans include
implementing wavelet transform algorithms on-board of upcoming CubeSat missions (e.g. TwinCube).
Due to CubeSat restrictions regarding computational power, energy and communication bandwidth,
algorithms using wavelet transform might find good application in this field. Although our wavelet
processing toolkit is developed for particular use on-board of electromagnetic plasma waves analyzers,
whole toolset might be applied to other types of scientific measurements where time-frequency analysis
and lossy data compression is required.
23
On feasibility of Global Radio Frequency Interference measurement with
CubeSats
J. Praks1, M. Vaaja1, J. Seppänen1, R. Modrzewski1, A. Hakkarainen1, S. Ben Cheikh1 and J. Lahtinen2
1
Aalto University, School of Electrical Engineering, Helsinki, Finland
2
Harp Technologies Ltd, Espoo, Finland
In this paper we discuss the idea of using CubeSats for Radio Frequency Interference (RFI) mapping on a
global scale, to provide up-to-date information for better satellite mission design.
The available radio frequency spectrum is a finite natural resource which is utilized more extensively
every year. Among other areas, radio frequency spectrum plays a key role in space technology and
microwave remote sensing, as it provides means for communication and sensing over large distances.
Despite strict international coordination, increasing deliberate or accidental transmissions on unwanted
frequencies create a problem for remote sensing satellites and satellite communication. This problem
needs attention on the global level. A good example of RFI-caused difficulty for an Earth Observation
instrument is the SMOS (Soil Moisture and Ocean Salinity, European Space Agency) mission which is
severely hampered by radio transmissions operating illegally at the protected frequency band of 1.4 GHz
all over the world. In the future, potential RFI problems should be better accounted for from the early
planning phase of missions. For this, the global RFI statistics and preferably global maps are required.
CubeSat provides cost-effective means for quick development of short precursor or technology
demonstration missions in LEO orbit. We propose that the platform can also be utilized for global RFI
mapping. However, size limitations of the platform set limits on the applicable frequency range,
radiometric accuracy and spatial resolution. In this work we provide trade-off calculations to identify the
limiting constraints and frequency ranges where RFI can cause problems for future missions. We show
that a precursor mission with CubeSats can provide valuable information for a larger mission to tackle
the RFI problem in the early design phase and provide information on RFI at the global scale.
24
Doing Science with University Cubesats: Present and Future
T. Moretto Jorgensen
National Science Foundation, Arlington, Virginia, USA
When the US National Science Foundation (NSF) began exploring the use of cubesats to conduct
space weather research in 2007 few people believed the miniature satellites would prove to be
a useful scientific tool. However, during the last five years, the NSF cubesat program has seen
the highly successful implementation of creative and innovative missions that carry out
important science experiments. Currently, the program supports 11 projects and has had 6
missions operating in space. The assortment of scientific investigations being pursued and
proposed spans all across solar, space physics, space weather, and atmospheric research.
Already, several projects in the program have delivered first-of-their-kind observations and
findings that have formed the basis for high profile engineering and science publications.
Inarguably, the results from the program have now established beyond doubt the scientific
value of cubesats and have proven them as a viable option for space missions that should be
taken seriously. Based on examples and lessons learned from current projects the presentation
will document and explore the prolific scientific promise of CubeSat missions.
25
The Micro Solar-Flare Apparatus (MiSolFA)
Diego Casadei
Fachhochschule Nordwestschweiz (FHNW), Bahnhofstrasse 6, 5210 Windisch, Switzerland
Solar flares are the most powerful events in the solar system and the brightest sources of X-rays, often
associated with emission of particles reaching the Earth and causing geomagnetic storms, giving
problems to communication, airplanes and even black-outs. X-rays emitted by accelerated electrons are
the most direct probe of solar flare phenomena. The Micro Solar-Flare Apparatus (MiSolFA) is a
proposed compact X-ray detector which will address the two biggest issues in solar flare modeling.
Dynamic range limitations prevent simultaneous spectroscopy with a single instrument of all X-ray
emitting regions of a flare. In addition, most X-ray observations so far are inconsistent with the high
anisotropy predicted by the models usually adopted for solar flares. Operated at the same time as the
STIX instrument of the ESA Solar Orbiter mission, at the next solar maximum (2020), they will have the
unique opportunity to look at the same flare from two different directions: Solar Orbiter gets very close
to the Sun with significant orbital inclination; MiSolFA is in a near-Earth orbit. To solve the crosscalibration problems affecting all previous attempts to combine data from different satellites, MiSolFA
will adopt the same photon detectors as STIX, precisely quantifying the anisotropy of the X-ray emission
for the first time. By selecting flares whose footprints (the brightest X-ray sources, at the chromosphere)
are occulted by the solar limb for one of the two detectors, the other will be able to study the much
fainter coronal emission, obtaining for the first time simultaneous observations of all interesting regions.
MiSolFA shall operate on board of a very small satellite, with several launch opportunities, and will rely
on moiré imaging techniques.
26
Mapping the radiation environment in Equatorial Low Earth Orbit with CubeSats
E. Del Monte1, M. Feroci1, Guido Parissenti2 on behalf of a larger collaboration
1
Istituto Nazionale di Astrofisica (INAF) - Istituto di Astrofisica e Planetologia Spaziali (IAPS), Roma (Italy)
2
Department of Energy - Politecnico di Milano, Milano (Italy)
Satellites in orbit around the Earth operate in a harsh particle environment. Thanks to past and recent
experiments, datasets and models of the proton and electron fluxes are available but with uneven
coverage, leaving some regions of interest for science-oriented satellites unexplored. Few data are
available about the radiation environment in Equatorial Low-Earth Orbits, with inclination below about 20
deg, where a number of astronomy missions require to operate. As an example, a low-energy and highly
directional population of protons have been recently discovered.
The currently available database leaves large uncertainties when developing science missions for these
orbits, with respect to both radiation damage and experiment background issues.
In the last years, compact, light and standard yet adequately sensitive particle monitors for satellite-borne
applications have been developed. CubeSats require low resources (in terms of mass and power) and are
relatively inexpensive, thus perfectly suited to carry in orbit such particle monitors and map the radiation
environment in the orbits of interest.
In this presentation we analyse the CubeSat mission profiles required to collect the required data about the
radiation environment in Equatorial Low-Earth Orbits.
27
From the low cost to the high benefits
The outreach of small satellites instrumentation
Marina Diaz Michelena
National Institute for Aerospace Technology (INTA)
Madrid, Spain
One of the technological challenges of the present century is to place laboratories and
observatories with global coverage in the Space. Constellations of small satellites are potential
platforms for such worldwide observatories as long as the payloads can stand the extreme
conditions of the Space: wide temperature ranges, vacuum, and radiation, and have low mass
and power consumption. But this idea is only feasible if the cost of the payloads is not
excessive.
Since the 1990s, motivated by the decrease in budgets in most of the world’s space agencies,
the use of Commercial Off the Shelf (COTS) has provided miniature devices with higher
functionality, lower power consumption, and lower cost than traditional high reliability (hi-rel)
and radiation hardened (rad-hard) components often used for flight.
At the National Institute of Aerospace Technology (INTA) of Spain, we developed and qualified
six COTS-based compact miniaturized magnetometers built with different magnetic
technologies: Anisotropic MagnetoResistance (AMR), Giant MagnetoResistance (GMR) and
MagnetoImpedance (MI).
These instruments are currently on several quasipolar Low Earth Orbits (LEO) onboard INTA
platforms:



NANOSAT-01 (19 kg mass launched in 2004)
NANOSAT-1B (23.9 Kg mass launched in 2009)
OPTOS (3.8 kg mass launched in 2013)
providing useful information for multiple purposes (Attitude Control System, COTS degradation
monitoring, Space weather, etc).
After more than ten years of experience we can state that the use of COTS in small plaftorms is
very profitable.
In this work we explain the resulting outreach of these developments, covering the
implementation of these miniaturized magnetometers on drones for geophysical prospections
to the planetary exploration on board landers and rovers.
28
A CubeSat for X-ray Polarimetry in Astrophysics
P. Soffitta1, E. Del Monte1, S. Fabiani2, F. Muleri1 and Guido Parissenti3
1
Istituto di Astrofisica e Planetologia Spaziali/INAF, Rome, Italy
2
INFN Sezione di Trieste, Padriciano (Trieste), Italy
3
Department of Energy - Politecnico di Milano, Milano, Italy
While the sun was the first non-terrestrial X-ray source discovered back in 1948, many aspects of solar
physics needs to be understood. One of this is the high energy emission following magnetic
reconnection events in the solar corona generating solar flares. In particular the expected X-ray
polarization from the impact of beamed electrons in the chromosphere must still be detected. Only
upper limit are available so-far derived from large instruments like RHESSI and SPR-N on Coronas F.
Celestial non solar X-ray sources first discovered in 1962 are also expected to be polarized in X-ray at
least of some extent providing information on the emission mechanisms, geometries and answering
questions of fundamental physics. Those include the Crab Nebula, for which the only positive
measurement results so far back in the ’70 and eventually micro-quasars harbouring a black-hole
surrounded by an accretion disk.
One characteristics of a polarization measurement in X-rays is that it requires a large amount of photons
to measure the few % expected. Sufficient amount of photons can be collected in case of a bright source
like the Sun or with very long integration time as in case of those much fainter celestial non-solar X-ray
sources.
For more than ten years, we exploits efficiently the photoelectric effect in gas to derive the polarization
of the incoming X-ray beam with a small, light and low-power device, called Gas Pixel Detector, based
on the use of a finely pixelated ASIC-CMOS chip that collects the charge amplified by a Gas Electron
Multiplier.
In this talk we present how such device allows for designing instruments capable of sensitive
measurements and meaningful results in the frame of the resources available by the class of nanosatellite CubeSat.
29
Evaluation of the performances of a spectro imaging detector (CALISTE) to be
embedded on board a nanosatellite for Solar Flares studies
H. Triou1, A. Meuris1, O. Limousin1, L. Gosset2 et al.
1
CEA Saclay, DSM/IRFU/service d'Astrophysique, 91191, Gif-sur-Yvette, France
2
EIDD, Université Paris Diderot, Paris, France
Solar flares (with CMEs) are the most violent phenomena in the solar system. Although general
characteristics of solar flare can be derived, each solar flare has unique characteristics: different
spectral indices for energy spectra can be found and the photon flux can be more or less important.
After an analysis of solar flares in X-rays we detail the constraints associated with the use of for the
CALISTE detector, which is an innovative and miniature sensor developed by CEA, for solar flares
studies. We present an analytical model that we realized in order to simulate the response of the
instrument. Many parameters have been analyzed to optimize the detector that will be embedded on
board a Nanosatellite (CubeSat type). The goal is to define a perfect counter and spectrometer which
means that the detector has to count the number of photons and determine their energies in order
to reconstruct the photon spectrum of the incident solar flare. This sensor is composed of 256 pixels
so that it is possible to switch on or off some of them according to the photons flux, in order not to
saturate electronics of the detector. A low level threshold can also be adjusted according to the
incident photons flux. Another important point will be the question of adding filters in front of the
sensor or not. The analysis of the response of the detector when varying these parameters allowed us
determines the number of counts that the detector’s electronics will be able to analyze perfectly.
Moreover, knowing the reduction of the number of counts obtained, we will be able to deduce the
incident photons spectrum in order to study it.
Finally we propose an optimal configuration of sensors which will be integrated to the nanosatellite
to observe the largest number of solar flares of different intensities as possible in order to study in a
quantitative way the solar flares phenomenon. In addition, the constraints associated with the
accommodation of our detectors on a nanosatellite are presented (type of orbit, volume of science
data to be downloaded via telemetry, power, mass, temperature requirements and so on …) and a
nanosatellite configuration (6U CubeSat) is proposed to meet these requirements.
30
Micropropulsion Systems
31
A Highly Miniaturized uPPT Thruster for Attitude-orbit Control
Junquan Li1, Steve Greenland1, Mark Post2, Michele Coletti3
1
2
Clyde Space Ltd. West of Scotland Science Park, Glasgow G20 0SP UK
Mars Space Ltd, Unit 61, Basepoint B.C., Andersons Road, S014 5FE, Southampton, UK
3
University of Strathclyde, 16 Richmond St, Glasgow, G1 1XQ, UK.
The successful miniaturization of spacecraft subsystems makes nanosatellites suitable candidates for
many scientific missions, and several miniaturized electric propulsion systems on CubeSats have
been studied. A Cubesat employing a Pulsed Plasma Thruster module is currently under
development by Clyde Space Ltd, Mars Space Ltd and the University of Southampton under contract
to ESA and is now entering final qualification testing. Pulsed Plasma Thrusters have been used in
satellite attitude control studies of nadir pointing, and satellite rendezvous, docking and formation
flying. This paper considers the use of the PPTCUP Pulsed Plasma Thruster in a range of different
operational cases that are considered key to the success of many future nanosatellite missions,
including moderate low thrust orbit inclination and altitude changes, deorbiting a satellite at the end
of life, maintaining “a string of pearls” constellation, on-orbit servicing and inspection of a mother
satellite, and life extension of an ISS-deployed nanosatellite. PPTCUP uses a very fast electric
discharge to ablate a tiny amount of a solid propellant bar, which is then ionized and expelled at high
velocity to generate thrust. To generate such a discharge, the thruster charges a high performance
capacitor bank for up to a second before discharging it in few microseconds. Based on a current
generation system with a mean power input of 2 W, a 40 uN thrust pulse (with a 20 N peak force)
can be achieved 1 million times in succession. In order to evaluate the PPTCUP thruster
performance, a simulation model has been developed and applied to the identified scenarios,
leading to suggestions for baseline control laws suited for the mission. Where necessary, refinement
of the design point has been performed to provide a roadmap for future PPTCUP development. We
assume a nanosatellite that has a maximum mass of 4 kg at 320-380 km. Drag force and yearly
velocity increments are estimated as functions of altitude, drag cross-section and solar activity.
Assuming an analysis of a 3U Cubesat frame with PPTCUP, a control approach based on advanced
control methods will be used to handle different operational cases. This work finds immediate
application for the PPTCUP for challenges such as drag compensation, de-orbiting, and rendezvous
and docking manoeuvres, and makes recommendations for future targeting of performance to
further improve the capability of PPTCUP equipped satellites. In particular it is noted that advanced
control design would improve the coverage of this class of thruster.
32
Solid Cool Gas Micro Propulsion System for CubeSat
Peng Zhu, Xiang Zhang, Ruiqi Shen, Yinghua Ye, Zhenhua Liang
Nanjing University of Science and Technology, Nanjing, China
This article studies the solid cool gas micro propulsion system suitable for CubeSats. The system includes
the gas generators, pressure sensor, valve and nozzle. Among them, the gas generators are composed of
igniter, gas generating agents and filter. The characterization of gas generator indicates that 100g gas
generating agents can produce approximately 66 normal liters of pure gas, and the temperature of the
gas squirting from the gas generator is about 53°C. The performance of solid cool gas propulsion system
was characterized by using ballistic pendulum. Results showed that the micro propulsion system is
suitable for the attitude control and orbit transfer of CubeSats.
33
Solid propellant micro-thruster array for CubeSat
Xuhui Liu, Yanming Wei, Jun Chen, Yan Shen, Xudong Wang, Jun Long
Institute of Control Engineering, Beijing, China
Solid propellant micro-thruster array has attracted widespread attention since the late 1990s. It can
provide micro impulses and impulse moments for precise orbit correction and attitude adjustment.
What’s more, it can provide combination among single thrusters for wide-range maneuverability. In
order to apply solid propellant micro-thruster array to CubeSats, allocation algorithm, ignition system,
design and testing of array’s principle prototype and space mission analysis, have been mainly studied in
this paper. Correlation functions of thruster impulse and moment distribution in CubeSat coordinate
system were established, that formed the mathematical model of thruster allocation. Then,
performance features of solid propellant micro-thruster array were analyzed and the dynamical models
of attitude and orbit control were established, respectively.
34
Development Status of an Open Capillary Pulsed Plasma Thruster with
Non-Volatile Liquid Propellant
E. Remírez(1), R. Martín(1), S. Barral(2), J. Kurzyna(2), A. Szelecka(2), H. Rachubinski(2), J. Miedzik(2), P. Ortiz(3), J.
Alonso(3), S. Botinelli(4), Y. Mabillard(4), P. Rangsten(5), A. Zaldivar(5), C.R. Koppel(6)
(1)
(2)
JMP Ingenieros, 26371 Sotés (La Rioja) Spain
Institute of Plasma Physics and Laser Microfusion (IPPLM), 01497 Warsaw, Poland
(3)
Najera Aerospace (NASP), 26371 Sotés (La Rioja) Spain
(4)
Mecartex, 6933 Muzzano, Switzerland
(5)
Nanospace, Uppsala Science Park, SE-751 83 Uppsala, Sweden
(6)
KopooS Consulting Ind., 75008 Paris, France
The desire to reduce development and launcher costs and the narrow focus of many payloads has in recent
years driven the development of very small platforms in the kg range. Such spacecrafts have benefited from
the wider availability of enabling technologies (micro/nano-fabrication), but remain hindered by the lack of
sufficiently compact and lightweight micro-propulsion systems. Low thrust propulsion systems have at the
same time also become a critical component in a number of scientific missions that require fine positioning,
such as space-based telescope interferometers, imaging arrays and formation flying missions . Due to their
simplicity and scalability, Pulsed Plasma Thrusters (PPT) are increasingly considered for small delta-V
missions on nano-spacecrafts.
The goal of the L-μPPT project is the development and assessment of a novel PPT technology based on an
open capillary design and on a non-volatile liquid propellant, which in comparison to conventional solidpropellant (PTFE) PPTs is expected to offer significantly larger total impulse and lower impulse bit variability
throughout the thruster lifetime. Another intrinsic advantage of liquid propellant is the possibility to
balance propellant requirements between several thrusters with a common tank, which in practice is
expected to enable to a twofold increase in propellant utilization.
The many issues associated with the use of water as liquid propellant and current lack of compelling
solution to address them has motivated the L-μPPT project to adopt a non-volatile liquid that can be easily
used and stored over a wide temperature range.
A novel type of Pulsed Plasma Thruster (PPT) based on an open capillary design and on a non-volatile liquid
propellant is currently under development within the Liquid Micro Pulsed Plasma Thruster FP7 project (LμPPT). Functional results from first prototype testing provide an Ibit 25 µNs @ 700V and 46 µNs @ 1000V,
presenting a Ibit/E ratio beyond 40 µN/W. Its design is expected to improve over PTFE-based PPTs by
providing significant increase in total impulse, increased propellant utilization, lower impulse bit variability
and the possibility to balance propellant requirements between several thrusters with a common tank.
35
CubeSat Launchers and Deployers
36
Reliable Low-cost Access To Space For CubeSat Size Payloads
P. C. Steimle1, C. Kuehnel2 and R. Pournelle3
1
2
Airbus DS GmbH, Bremen, Germany
Astrium North America Inc., Houston, Texas, United States of America
3
NanoRacks LLC, Houston, Texas, United States of America
In recent years the commercial use of space has seen significant dynamics, especially in the field of
small satellites, as more and more countries get involved in space business. But the growing small
payload market is unlikely supported by the launcher systems and mission concepts currently
available which creates a bottleneck situation that could slow down market growth significantly. One
important implication of the small payload dilemma is the difficulty to access the optimal orbit for
the mission at acceptable system and launch cost. The common practice of sharing the launch with
other small satellites often leads to a compromise between the cost effective launch and the perfect
satisfaction of payload mission requirements. This is especially true for the majority of technology
demonstration missions. The commercial utilisation of the International Space Station (ISS) for smallsize payloads driven by the U.S. company NanoRacks LLC is one opportunity to overcome this small
payload dilemma. Two concepts are offered to the market: One is the External Payload Platform
(EPP) designed and manufactured by Astrium North America and Airbus Defence and Space for
unpressurised CubeSat size payloads which will be launched to the ISS and installed on the Japanese
Experiment Module External Facility (JEM-EF) by the end of 2014. The other is the well-known
CubeSat Dispenser (CSD) which deploys CubeSats from the ISS on a regular basis. The two concepts
and their opportunities and constraints are presented focussing on the needs of small payloads. The
EPP design allows the fully robotic installation and operation of payloads. In the nominal mission
scenario payload items are installed not later than one year after the signature of the contract, stay
in operation outside of the International Space Station for 15 weeks, and can be returned to the
customer thereafter. Payload items are transported among the pressurised cargo usually delivered
to the station with various supply vehicles. Due to the high frequency of lights and the flexibility of
the vehicle manifests the risk of a delay in the payload readiness can be mitigated by delaying to the
next flight opportunity. The mission is extra-ordinarily fast and of low cost in comparison to
traditional activities conducted on-board the ISS and can fit into short-term funding cycles available
on national and multi-national levels. This fast turnaround can also help payload developers
maintain support for their programs by providing tangible in-space testing. The standard payload
size is a multiple of a 4U CubeSat. Every payload can extensively use all ISS resources required: mass
is not limited, power only limited by the payload heat radiation capability, the data link is a USB 2.0
standard bus enabling a real-time and private data link to the payload operator's work station. The
new EPP transforms the station into a true laboratory in space with the capability to support
research and development in various fields as well as in-orbit demonstration and verification. The
CSD also provides reliable access to space through regular flights to ISS. The CSD can deploy satellites
ranging from 1U to 6U including the 2Ux3U form factor. Currently, 2-3 CubeSat missions, deploying
30 – 40 per mission are performed each year.
The EPP and the CSD both are operated in the frame of end-to-end services dedicated to be of low
cost and while introducing a new level of technical and programmatic reliability to small payload
missions to support current market dynamics.
37
On-Orbit Optimization of CubeSat Launch Opportunities
using an Orbital Maneuvering Vehicle
M. Stender1, J. Maly2, C. Loghry3, C. Pearson1, E. Anderson2
1
Moog, Inc. Space and Defense, Advanced Missions and Science, Golden, CO, USA
2
Moog CSA, Mountain View, CA, USA
3
Moog In-Space Propulsion, Chatsworth, CA, USA
As the capability of CubeSats increases, one limiting factor remains – non-ideal drop-off orbits. Although
secondary payload opportunities are becoming more available, CubeSats are still accepting less than
optimal orbits in order to access space. In the case of constellations, commercial and government
entities are now recognizing their potential, but still need a way to quickly and economically adjust the
orbits of these valuable secondary payloads.
The mobility of individual CubeSats and smallsats is inherently constrained by the limited volume
available for a propulsion system. However, further factors also limit them, such as the cost of a
miniaturized system (assuming technology is available), the perceived hazards, and the time required for
a large delta-V maneuver (in the case of EP). By utilizing a single orbital maneuvering vehicle (OMV), a
number of these challenges can be overcome and the burden of a propulsion system is not imposed on
a CubeSat. Moog’s family of OMVs is uniquely suited to this task due to a range of modular options that
build on the proven ESPA structure.
Previous missions, such as LCROSS and the upcoming launch of DSX, have taken advantage of such
modularity to easily integrate avionics, solar panels, and even the upper stage of a launch vehicle. New
designs for CubeSat and small sat deployment mechanisms also provide secondary payloads with a
wider range of options for staggered deployment in one or multiple orbits. An example OMV
configuration and a representative CubeSat constellation deployment timeline will be presented to
illustrate how a secondary payload can share a launch and still achieve their desired mission parameters.
38
Swiss Space Systems Innovative Launch Concept for Small Satellites
B. Deper, P. Jaussi and B. Vuitel
Swiss Space Systems Holding SA, Payerne, Switzerland
Swiss Space Systems (S3) is a young company with the objective of providing a reusable, flexible,
safe and efficient launch system for small satellites. The Swiss Space Systems launch model
comprises of an Airbus A300 aircraft (first stage), a sub-orbital shuttle “SOAR” (2nd stage) and the
third stage spacecraft with the satellites inside. The S3 launch concept uses the Airbus A300, an
aircraft already certified for zero gravity flights, to take the SOAR shuttle up to 10 km on its back; the
shuttle will be launched from there. The shuttle will then ascend up to an altitude of 80km, the
height at which the upper stage will be launched in order to put the satellites into orbit. Once this
operation is completed, the shuttle will return to earth by gliding towards its launch airport, where it
will be taken care of by the maintenance teams who will prepare it for a new launch. The system
developed by S3 has many safety advantages: the launch can be terminated and the shuttle can
return to Earth at any time during the process. With first and second stages that are regularly reused
and a fuel consumption that is much lower than at present, Swiss Space Systems will be able to offer
satellite launches at approximately four times less than current market prices. Any large airport is
suitable to act as a “space port” and there is no need to erect a launch site from scratch.
The company mission is to give access to space. S3 wants to make space accessible through fast and
recurrent access opportunities facilitating particularly science and in-orbit delivery. The company
objective is to develop, manufacture, certify and operate unmanned suborbital space planes for
small satellite deployment. The range of satellites to be launched goes up to 250 kg small satellites.
The start of the test flights is planned for 2017, the first commercial flights in 2018. In order to
achieve this goal, S3 relies on the support of a worldwide network of internationally renowned
partners and advisors, who all support S3 and trust our vision.
39
Technology Demonstration
on CubeSats
40
Space Object Detection on a CubeSat platform
M. Tetlow1, P. Veitch2 and T. Chin2
1
2
Inovor Technologies, Adelaide, Australia
The University of Adelaide, Adelaide, Australia
Space Situational Awareness (SSA) is of key interest to Australia, as stated in defence and civilian policy
documents. Considerable investment has been made in ground based systems to detect space objects.
These systems can detect and track objects with high accuracy. As with all ground based platforms,
depending on the orbit of the unknown object, considerable time may elapse between observation
opportunities, and in some cases, the object may never be observable from a particular location on the
surface of the Earth. A low cost space based detection system of ~5 nano satellites would complement
these ground based systems by providing much wider detection coverage.
The concept of operations (CONOPS) is for a small evenly-distributed constellation of ~5 CubeSats to
conducting SSA observations. The SSAsat will capture an image and compare it to the star chart
onboard. Using a subset of identified stars, it will estimate the attitude of the space craft and then
remove all known stars from the image, based on the star chart. The remaining objects will then be
compared to the catalog of known space objects in the NORAD catalogue. If any objects are neither in
the star chart nor the NORAD catalogue, they will be identified as unknown. The SSAsat will then
attempt to capture more images of the object to estimate its orbit. Information about unknown objects
will be handed over to other CubeSats in the constellation to provide the best chance to tracking. With
over 80% of the space investment being in the GEO and MEO belts, these will be the target orbits.
This paper presents a nano-satellite based imaging system that can conduct space situational awareness
monitoring in the Medium Earth Orbit and Geostationary Earth Obit belts from Low Earth Orbit. A
systems analysis and configuration design of the nano-satellite is presented, showing how the imager
will be integrated into the nano-satellite. Proposed COTS and custom designed hardware is presented
along with relevant budgets, showing the feasibility of the design. A robust algorithm for detecting false
stars is also presented to support attitude estimation as well as detecting unknown space objects.
41
The Multi-Payload Satellite Service (MPS) for in-orbit technology evaluation
L. León1, J. M. Quero2, J. M. Moreno1
1
2
Solar MEMS Technologies S.L., Seville (Spain)
Department of Electronic Engineering, University of Seville (Spain)
According to ESA criteria, in-flight performance is the highest level of qualification for space products.
For this reason, there is an increasing demand for in-flight services to accommodate new devices and
systems for their evaluation and validation. The Multi-Payload Satellite (MPS) is a service, offered by
Solar MEMS, to evaluate new technologies under a low-cost CubeSat mission. This service provides a
complete support for every mission step, leaving to the technology developers only the design and
manufacturing of their own payloads. MPS includes the CubeSat platform assembly, payload integration,
operational tests, launch, in-orbit evaluation and test reports. The mission budget is shared between
several onboard payloads, thus reducing the service cost for each one under a similar idea than the
multi-project wafer (MPW) in microelectronics. MPS also includes consulting services and design
assistance if required, in order to offer the flight opportunities to any manufacturer, even without
previous knowledge about a space mission.
The suitability of CubeSats as platforms for payloads experimentation is proven with the first mission
CEPHEUS. CEPHEUS is a 3U CubeSat with five different payloads for in-flight experimentation, developed
by various companies and full integrated by Solar MEMS and the University of Seville. The payloads
consist on a fuel cell with an interface control unit, a miniaturized Star Tracker, a high quality radio
transceiver, a high accurate MEMS sun sensor and an advanced attitude control software. During this
mission, payload integration procedures have been standardized, optimizing the mission in terms of
design time and costs. This project is the starting point of the MPS service and is scheduled for launch in
the first half of 2016.
MPS offers the opportunity to reach the highest competitiveness level for a CubeSat product, which is
the in-flight validated. The integration of payloads in a proven quality platform, based in a rapid
development approach, provides a flexible solution to gain heritage and dramatically reduces the
mission time, cost and complexity, opening new opportunities and markets.
42
SamSat–QB50 nanosatellite.
Burning wires mechanism for antenna system and aerodynamic stabilizer
E. Ustiugov1, A. Nikitin1, S. Shafran1
1
Samara State Aerospace University, Samara, Russia
SamSat-QB50 CubeSat contains two deployment systems: antenna system and aerodynamic stabilizer.
Antenna deployment system (AntS) is part of communication system based on COTS transceiver. AntS
has four spring rods with are produce the force for deploy. This spring rods are twisted like helix inside
of the special space in AntS. This space are blocked by the door with are locked by the burning wire.
Aerodynamic stabilizer (AeroS) is the part of ADCS based on using aerodynamic force and magnetic field.
AeroS has a bottom plate, rods and balloon with are produce the force for deploy. Before
transformation bottom plate is locked by the burning wire.
Both of two systems deploy separately and after command from OBC. It means two controllers are
needed for each if this systems, but because the designs are in-house, only one controller is used.
Commonly AntS controller is designed for four controllers and four burning mechanism. Four is staff and
four is reserved but the controller can handle more than eight commands. We transfer functions of
AeroS controller to AntS controller. Command to deploy AntS arrives at AntS controller and switches on
the burning wire mechanism. Command to deploy AeroS arrives at AntS controller and is transferred to
the burning wire mechanism of AeroS.
This principle allows designing simpler deployment systems with less number of components and better
reliability. Laboratory tests show the normal and stable functioning of deployment mechanisms of both
systems.
43
Testing of low-cost GNSS receivers for CubeSat orbit and attitude determination
C. Hollenstein1, B. Männel1, E. Serantoni1, L. Scherer1, M. Rothacher1 and Ph. Kehl2
1
Geodesy and Geodynamics Lab, ETH Zurich, Switzerland
2
u-blox AG, Thalwil, Switzerland
In the cooperative Swiss CubeSat project ‘CubETH’ - involving ETH Zurich, the Swiss Space Center,
several universities of Applied Sciences, and Swiss companies - the main science goals are precise orbit
determination and attitude determination of a (1-unit) CubeSat using single-frequency GNSS receivers.
For this purpose, CubETH will be equipped with u-blox multi-GNSS receivers - low-cost, single-frequency
COTS receivers for embedded solutions, providing navigation solutions as well as raw code and phase
measurements, characterized by good performance and very small size, weight and power consumption
and, therefore, predestined for CubeSat missions. However, they are not space-qualified. Therefore, in
the first phase of the project, numerous tests have been performed in order to study the behaviour of
the receivers in the intended space environment and to evaluate their usability for space applications.
In this presentation we focus on the receiver tests carried out - in particular radiation, temperature,
vacuum and GNSS simulator tests - shortly describing test setups and procedures and presenting current
results and conclusions drawn from these tests. Apart from types and frequencies of radiation effects
that have to be expected in the orbit, the results of the radiation tests show a remarkable autonomous
error detection performance and a generally good recovering capability of the receivers. In the
temperature vacuum cycling tests, we were mainly interested in the physical resistance of the nonspace-qualified receiver components against vacuum and the exposure to extreme temperature
conditions as well as the behaviour of the internal clock under these extreme conditions and its effects
on the measuring performance of the receiver. GNSS simulator tests revealed valuable information on
the receiver’s performance in tracking and measuring GNSS under space conditions and its impact on
the scientific tasks. The ‘untuned’ internal u-blox navigation solutions (code solution) of a simulated LEO
orbit with a height of 450 km revealed accuracies of about 3-4 m in position and <10 cm/s in velocity.
The results of the tests carried out up to now support the conclusion that the u-blox receivers - although
not fully resistant to effects due to space environment - are usable for space applications in LEO orbits
as planned in the project, if the concept includes latch-up protection and redundant receivers.
44
QARMAN: As an Atmospheric Entry Experiment on CubeSat Platform
M.E. Umit, V. Van der Haegen, G. Bailet, I. Sakraker, T. Scholz, P. Testani
von Karman Institute for Fluid Dynamics, Rhode-Saint-Genèse (Brussels), Belgium
QARMAN, QubeSat for Aerothermodynamic Research and Measurements on AblatioN, is a triple unit
(3U) CubeSat that will perform an experiment on Earth atmospheric entry. QARMAN has three payloads,
which will operate on different time slots of the mission.
The first payload “Semi-controlled differential-drag-based manoeuvres” experiment will be conducted
after commissioning and de-tumbling phase. The aim is to control the surface exposed to the residual
atmosphere, changing the magnitude of the atmospheric drag and therefore creating a (differential)
force, between one spacecraft (chaser) and either another spacecraft or a desired target point.
The main QARMAN payload is the usage of a CubeSat platform as “Atmospheric Entry Demonstrator”.
Spacecraft descending towards a planet with an atmosphere experience very harsh environment as
extreme aerodynamic heating and exothermic chemical reactions occur due to the gas surface
interaction at hypersonic free stream velocities. Such vehicles have special shields to survive these harsh
conditions, so will QARMAN. If this mission is successful, different entry vehicle configurations can be
tested on board at very low costs for scientific exploration and qualification of future missions in order
to provide valuable real flight data.
To collect real flight data the challenging physics of atmospheric entry to be investigated are downselected to make scientifically valuable measurements respecting the constraints of CubeSat platforms.
Thermal Protection System (TPS) ablation, efficiency, and environment; attitude stability; rarefied flow
conditions; off stagnation temperature evolution and finally aerothermodynamic environment will be
measured on QARMAN using COTS spectrometer, photodiode, temperature, pressure sensors. The
feasibility study of an effective TPS that could fit within the external dimensions of a 3U standard
CubeSat is one of the challenging parts of this project. It has to manage the thermal environment until
the targeted altitude, by keeping the payload bay in a suitable temperature.
QARMAN mission aims to provide an Earth entry flight data set for a given entry trajectory. This requires
an accurate de-orbiting system for QARMAN to reach 7.7 km/s at 120 km altitude. Thus, the third
payload of QARMAN is called “Aerodynamic Stability and De-Orbiting System (AeroSDS)”. The AeroSDS
will demonstrate the feasibility of a passive system providing aerodynamic stability for a CubeSat below
350 km of altitude.
45
A Thermal Protection System for a Re-Entry CubeSat
P. Testani, M.E. Umit, V. Van der Haegen, T. Scholz, I. Sakraker, G. Baillet
von Karman Institute for Fluid Dynamics, Rhode-Saint-Genèse (Brussels), Belgium
QARMAN (Qubesat for Aerothermodynamic Research and Measurements on AblatioN) is a 3U CubeSat
designed to collect scientific data during re-entry to Earth atmosphere.
The thermal design of the QARMAN CubeSat, with special attention to the re-entry phase, is a major
topic. In fact during the atmospheric re-entry, the CubeSat will interact with the atmosphere at
hypersonic velocity and, due to aerodynamic heating and exothermic chemical reactions, it will face
temperatures which can go over 2000 K.
Protecting the CubeSat components from those heat fluxes is one of the most critical aspects of the
mission: designing a TPS capable to protect the satellite within the standard dimensions of a 3U
CubeSat, is a challenging and delicate task. After a preliminary study, the QARMAN team efforts were
oriented to protect only those components necessary to complete the re-entry phase of the mission,
designing a “Survival Unit” capable to keep the electronic components within the operative limits for the
entire re-entry phase. Nevertheless this design shall be thermally compatible with the orbital thermal
environment as well.
The design solution presented shows the implementation of the Survival Unit, where only the
components supposed to survive the re-entry phase are placed. For this design solution different
thermal analyses have been performed. The cases run embrace the hot and cold worst cases for orbital
phases and the worst case scenario for the re-entry phase.
46
A tool for nano-satellite functional verification: comparison between different
in-the-loop simulation configurations
L. Feruglio1, R. Mozzillo1, S.Corpino1 and F. Stesina1
1
Politecnico di Torino, IT
This paper describes the simulator technology and the verification campaign for the e-st@r CubeSats
family, developed at Politecnico di Torino. The satellites’ behavior has been investigated using a Model
and Simulation Based Approach. One of the critical issue in the verification and validation of any space
vehicle is the impossibility to fully test some features due to the particular and often un-reproducible
environment in which it will operate. Simulations result as one of the best means for testing space
system capabilities as it may help to overcome the abovementioned problem.
In order to perform different simulation configurations for e-st@r CubeSats, an in-house simulator
(named StarSim) has been developed. It is a unique infrastructure, modular and versatile, capable of
supporting any desired configuration of the system under test, ranging from full algorithm in the loop
simulations (AIL), and gradually inserting satellite hardware, until a complete hardware in the loop (HIL)
simulation is performed.
When a verification campaign is led on a real object, pure AIL computer based simulations (in which all
the equipment and mission conditions are reproduced by virtual models) are not sufficient to test the
actual software and hardware to a high degree of confidence since real systems can exhibit random and
unpredictable dynamics difficult to be perfectly modeled (i.e. communication delays, uncertainties, and
so on). For these reasons, Software In The Loop (SIL), Controller In The Loop (CIL) and HIL simulations
were planned. SIL simulations foresee that algorithms are written in the final programming language
and executed on ground hardware. In CIL simulations, the software runs on the flight processor while
other system’s element are still kept virtual. In HIL simulation, the real hardware (i.e. sensors, actuators,
and power sources) are included in the loop.
In this paper, after the details of the simulator architecture and its characteristics are described, an
exhaustive comparison between AIL and HIL simulations is presented, highlighting main differences and
singularities: similar trends of the sensible system’s variables are reached but not identical
performances (i.e. absolute and average pointing error and stability, attitude determination accuracy,
battery charging and discharging duration) arose analyzing the values. Moreover, it is demonstrated
how the technology here presented can effectively support and improve the verification and validation
activities for a nano-satellite, by increasing the confidence level on the mission objectives achievement.
47
Telecommunications, Ground Stations,
Ground Station Networks
48
Development of an Open Source Software Defined Radio
M. Wegerson1, J. Straub2, S. Noghanian1, R. Marsh2
1
Department of Electrical Engineering, University of North Dakota, USA
2
Department of Computer Science, University of North Dakota, USA
The Open Prototype for Educational Nano-satellites (OPEN) is a CubeSat design being developed by a
faculty-mentored student group at the University of North Dakota. Its primary goal is to create a lowcost, dependable satellite framework that can be utilized for a variety of missions. When completed,
this design will be available to universities around the world allowing them to focus on research and
development of the primary mission payload and less on designing the satellite’s supporting
subsystems. In aliment with this goal, a software-defined radio (SDR) unit is being implemented for the
satellite’s primary radio. This should facilitate greater versatility in space-ground and ground-space
communications.
The novelty of this SDR design is that it is based off of the use of the Raspberry Pi micro-computer as the
primary transmitter and a USB FM Radio receiver containing the Realtek RTL2832U chip as the primary
receiver. Through the use of open source software, the Raspberry Pi has the ability to transmit FM
signals ranging from 1 MHz to 250 MHz. In addition, a USB Radio receiver allows for reception of
frequencies from 24 MHz to 1766 MHz.
Current work to date includes working with the software that allows the Raspberry Pi to transmit and
working with GNU Radio to set up a working SDR receiver with the USB Dongle. Work began several
months ago with setting up the USB TV tuner for reception of FM radio signals. Although there was
some initial challenges with interfacing the USB device with GNU Radio, we were successful at receiving
basic FM signals in November 2013 and have since been working towards improving the GNU Radio flow
chart with the addition of several dynamic filters and data management systems.
Work towards the transmitter has focused mainly on developing the Raspberry Pi into a functioning
transmitter. For this we are using two open-source programs to modulate and transmit the data:
minimodem for modulation and PiFM for transmission. Minimodem is a program that converts binary
data into a modulated audio signal using AFSK (audio frequency shift keying) and outputs a .wav file.
This file is then in turn transmitted by PiFm via GPIO pin 4. After the audio file has been received, it is
decoded by minimodem back into its original format. Several issues we have faced is PiFm dependence
on the operating system. Slight changes and updates to the operating system that alter the way
programs interface with the Raspberry Pi sub-systems drastically affect the usability of PiFm. Currently,
we are using Arch Linux as the control OS and have been successful at transmitting data at 300 and 1200
baud.
Due to limitations on the Raspberry Pi’s transmitting abilities, the square wave is generated for the FM
output. To counter the odd harmonics that are generated, we have designed a low-pass filter on the
transmitter side to remove these unwanted transmissions. Currently, we have a filter tuned for 144.39
MHz but we are actively researching a method for a variable filter design to allow for a greater range of
transmittable frequencies. In addition, we also have been working on developing improved, low-cost
antenna and amplifier designs to improve the power of the transmitted signal.
49
CubETH COM subsystem
F. Belloni1, A. Ivanov1,L. Van Box Som1, G. Laupre1, V. Richoz1
1
École polytechnique fédérale de Lausanne, Lausanne, Switzerland
CubETH is a joint project of Swiss Federal Technical schools (ETHZ and EPFL) with the PI form ETHZ and
supported by Universities of Applied Science. The main objective is to build a satellite scheduled for
launch in 2016. The CubETH spacecraft will be capable of calculating its own position in space with
unprecedented precision thus paving the way for nano-satellite constellations with inter-satellite
communication capabilities.
Compared to the Swisscube, the first Swiss satellite, the amount of data generated is considerably
increased and new fastest communication boards have been designed. The old design approach using
analog discrete components has been replaced with digital IC design. This allowed to simplify the PCB
design thanks to the reduced RF electronics, to reduce the PCB used surface thanks to the smaller
amount of components. This can permit to design a fully redundant communication system on a single
CubeSat PCB board.
In addition to improved reliability, we implement error correction. For uplink a BCH decoding and
downlink Reed–Solomon and convolutional encoding. Due to the low processing power available a
special attention to this task is required.
The goal is to have a hardware that can communicate at data rate higher than 9.6 kbps, with the power
consumption in Rx mode lower than 100 mW and a with a sensibility better than -110 dBm.
50
Enhancements of antenna control and an error correction scheme
for the CubETH Ground Station
S. Kaufmann1, R. Müller1, M. Joss1, M. Klaper2
1
Lucerne University of Applied Sciences and Arts (HSLU),
Dpt. Electrical Engineering, Horw (Lucerne), Switzerland
2
Lucerne University of Applied Sciences and Arts (HSLU),
Dpt. Computer Science, Horw (Lucerne), Switzerland
CubETH, a scientific pico-satellite, is currently under development. Its main mission goal is to measure its
position in space and exact attitude with a high degree of accuracy, using commercial off-the-shelf (COTS)
GNSS receivers. One of the tasks of HSLU is to implement the ground station.
As it is the case in almost every satellite project, the downlink is of high mission priority. This lies in the
nature of not having sufficient power onboard to transmit information at high power levels. Also there are
limited opportunities for high-gain antennas. Naturally, this results in receiving very low power levels down
on earth. The same is the case with CubETH. Thus, back on ground, the incoming signal needs to be received
with as much gain as possible. Therefore, four 22 element UHF Yagi antennas are stacked to get a high
antenna gain. However, with high gain, the antenna system needs to point exactly towards the angle of the
incoming signal. During the pass of the satellite, the antenna system needs to be continuously updated in
azimuth and elevation angles by means of closed-loop control. In order to generate an error signal for
controlling the movement of the antenna system, the concept of beam splitting and radar monopulse
technique is used. The downlink antenna array is excited out of phase so that the main beam of the
radiation pattern can be split. Combined with the sum signal of all four UHF antennas, a meaningful error
signal can be provided. The results show, that it is possible to operate the 435 MHz antenna movers in a
closed loop control.
On the other hand, a professional space link conforming to ECSS/CCSDS standards brings some important
advantages. The main one is the forward error correction feature. The CubETH space link features a binary
BCH block code for the uplink and a concatenated coding system for the downlink. The concatenated coding
system uses the Reed-Solomon block code as outer and a ¾-rate convolutional code as inner code. This well
performing and reliable coding system has evolved over time to its best. A good example of its convincing
capabilities can be given by looking at the voyager spacecrafts, which are yet the furthermost travelling
objects engineered by mankind. They use a similar coding system to the CubETH. The mentioned coding
system is implemented in the CubETH TMTC modem, which is a software only solution. Together with the
soundcard of a PC, it is possible to link to transceivers very universally and ham-friendly via an audio FSK
interface signal. By the presented two measures the net data rate can be improved.
51
SUSat Communications System
W.G. Cowley1, H. Soetiyono1, R. Luppino1, T. Kemp1, J. Kasparian1 and F. Ishola2
1
Institute for Telecommunications Research, University of South Australia, Mawson Lakes, Australia
2
International Space University (ISU), Strasbourg
SUSat is a two-unit CubeSat designed for the QB50 project by the University of Adelaide and the
University of South Australia. At UniSA, the Institute for Telecommunications Research (ITR) is
responsible for the SUSat communication system, including ground station. SUSat will use the 70 cm
and 2m amateur radio bands for communications with ground stations and other QB50 CubeSats.
This paper describes a custom communications payload design for SUSat. To meet demanding power
and mass constraints in the CubeSat the baseline design is based on the use of a COTS FSK/ MSK RF IC.
An additional power amplifier provides over 20dBm RF power. The design has several novel features,
including: the addition of a balanced modulator to allow BPSK transmissions, a new layer-2 protocol
called SKLEP to avoid the use of AX-25, channel coding options to reduce bit errors, plus flexible data
rates and transmit/ receive duty cycles, plus GPS-synchronised frame timing to improve efficiency. The
SUSat CubeSat also includes a custom module for the UHF and VHF antennas. The ground station will
use a software defined radio (SDR) approach.
The paper provides an overview of the SUSat communications system design and describes current
status and test results.
52
Antenna Subsytem of GAMALINK Platform
V. Akan, C. Dudak
TUBITAK Space Technologies Research Institute, Turkiye
Nowadays, there is a remarkable interest to CubeSats as they offer cheap and easy platforms for
research, technology demonstration, scientific and educational space applications. Usual functions like
attitude determination and control, uplink and downlink telecommunications, and power subsystem are
performed in a Cubesat.
GAMALINK is a platform that adapts the terrestrial communication and attitude determination
technologies into space applications based on Software Defined Radio (SDR) hardware. The main goal of
the platform is combining and integrating the miniaturized modules which will perform for mobile adhoc networking, attitude determination, GPS signal receiving, intersatellite and ground station
communication from the existing terrestrial technology. The entire design has been made taking into
account the CubeSat constraints and the space environment. On GAMALINK platform, there exists
different designs leading the developed modules, relating fields of RF electronics, antenna, acquisition
and signal processing. In order to save transmission power, beamforming will also be used by combining
RF electronics and antenna modules for communication.
Since most of CubeSats use amateur UHF band (about 437 MHz) for their communication, it provides
very limited bandwidth. To increase the communication bandwidth, ISM S-Band has been chosen for
high-speed data transmission and related communication. This communication is not only applicable
between CubeSat and ground station but also among CubeSats in a network. For beamforming
capability 3 S-band Antenna element array is used. Each S-Band Antenna has 24mm x 24mm x 3.2mm
maximum dimension to fit onto 2U surfaces. Moreover, for receiving GPS signal there is one antenna
whose dimension is 28mm x 28mm x 3.2mm on each lateral panel. Since major intended challenge of
GAMALINK is fitting the modules into Cubesat, miniaturization techniques have been utilized for S-Band
and GPS antennas.
53
A Simple Miniaturized Printed Antenna Adaptation for
CubeSats and Small Satellites
S. Kose, V. Akan, C. Dudak, E. Oncu
TUBITAK Space Technologies Research Institute, Turkiye
Today small satellites and CubeSats are very popular since they provide fast and cheap platforms for
technology demonstration, experimental research, educational purpose, etc. However there are some
restrictions like physical limitations. To overcome this difficulty subsytems and modules should be as
small as possible. This case is also valid for antennas which are initial/final elements of RF
communication systems. In the literature, there are five general antenna miniaturization techniques:
Geometrical Shaping (slot loading, bending and folding, meandering, and etc.), Material Loading (using
with dielectric and magnetic material), Lumped element loading, Optimization Methods (Genetic
Algorithm, Particle Swarm Optimization, and etc.), Miniaturization using artificial engineered
electromagnetic metamaterials. In this manuscript, a miniaturized microstrip antenna via geometrical
shaping has been analysed to use on CubeSats and Small Satellites.
The analyzed circularly polarized antenna has been simulated on a full-wave electromagnetic simulator
and then a prototype has been manufactured on Rogers TMM6 substrate with a relative permittivity of
6.0 and a thickness of 5.08mm and a loss tangent of 0.0023 and tested. The simulation and test results
are in good agreement. The achived antenna size is 30mm x 30mm. Measurements and simulations
have been realized on a 100mm x 100mm ground plane to model 1U CubeSat surface. 10-dB return loss
bandwidth and peak gain have been measured to be approximately 100 MHz and 7dBic at the operating
frequency through S-Band, respectively.
54
Development of an X-Band Transmitter for CubeSats
S. Palo1, D. O’Connor2, E. DeVito2, R. Kohnert2, S. Altunc3 and G. Crum3
1
Department of Aerospace Engineering Sciences, University of Colorado, Boulder, CO USA
Laboratory for Atmospheric and Space Physics, University of Colorado, Boulder, CO USA
3
Goddard Space Flight Center, NASA, Greenbelt, MD USA
2
CubeSats have developed rapidly over the past decade with the advent of a containerized deployer
system and ever increasing launch opportunities. These satellites have moved from an educational tool
to teach students about engineering challenges associated with satellite design, to systems that are
conducting cutting edge earth, space and solar science. Early variants of the CubeSat had limited
functionality and lacked sophisticated attitude control, deployable solar arrays and propulsion. This is
no longer the case and as CubeSats mature, such systems are becoming commercially available. The
result is a small satellite with sufficient power and pointing capabilities to support a high rate
communication system.
Communications systems have matured along with other CubeSat subsystems. Originally developed
from amateur radio systems, CubeSats have generally operated in the VHF and UHF bands at data rates
below 10kbps. More recently higher rate UHF systems have been developed, however these systems
require a large collecting area on the ground to close the communications link at 3Mbps. Efforts to
develop systems that operate with similar throughput at S-Band (2-4 GHz) and C-Band (4-8 GHz) have
also recently evolved. In this paper we outline an effort to develop a high rate CubeSat communication
system that is compatible with the NASA Near Earth Network and can be accommodated by a CubeSat.
The system will include a 200kbps S-Band receiver and a 12.5Mbps X-Band transmitter. This paper will
focus on our design approach and initial results associated with the 12.5Mbps X-Band transmitter.
55
Multi-objective optimization of a high gain, circularly polarized rectangular
antenna array in the Ka band for CubeSat class satellites
Alessandro Cuttin1, Livio Tenze2, Roberto Vescovo1
1
Università degli Studi di Trieste, Dipartimento di Ingegneria e Architettura
2
Esteco S.p.A., Area Science Park
Over the last decade, more than one hundred CubeSats have been launched, and this number will
double very soon. However, the increasing number of launched satellites is not yet followed by an
improvement in the communication technology they use: to date, the majority of orbiting CubeSats
transmits their data using frequencies allocated in the radio amateur service.
A reasonable explanation for this scenario is that most organizations face their very first experience in
the space sector and, therefore, prefer to use more reliable and simple solutions. Communication
systems featuring faster data rates are generally employed after the first mission. If this trend does not
change in favor of more advanced technologies, this class of satellites will not be considered a viable
solution for commercial purposes, and thus relegated to the role of educational tool or technology
demonstrator. The transition to higher frequencies, like those in the Ka band, will enable the design of
small and directive antennas, and will make faster data rates attainable -- up to 100 Mbps.
In the nanosatellites development, an antenna can be very demanding in terms of mass and real estate;
efficient radiators such as horn antennas or parabolic reflectors in most cases are not viable solutions
because of their mass and surface requirements. An ideal antenna should be thin, flat, small, light and
designed to be mounted on the outer surface of the spacecraft. In this context, a microstrip array
antenna is a convenient solution in terms of size, mass, mounting and cost.
The proposed antenna will be circularly polarized and will have a uniform polarization over a wider
bandwidth, respect to the bandwidth of the basic element.
In the design of this array, many parameters are subject to optimization: specifically, gain and
bandwidth shall be maximized, while the axial ratio and insertion losses shall be minimized. These
parameters can be optimized at every stage of the design: for the single patch, the sub-array, and the
final array.
In order to identify the best compromise possible between physical parameters and electromagnetic
performance, a multi-objective optimization approach will be adopted, using a CST MWS® parametric
model together with ESTECO’s modeFRONTIER® multi-objective optimization and process integration
tool. Numerical results will be illustrated.
56
KSAT light – a low cost ground station network for CubeSats
B.Eilertsen1, M.Krynitz1
1
Kongsberg Satellite Services AS, Tromsø, Norway
As Cubesats are becoming more and more potent, data rates from their payloads are increasing quickly.
This calls for using S-, X- or Ka-band transponders as on bigger satellites. KSAT, a leading supplier of
commercial ground station network services is currently bringing online a brand new network of small
aperture antennas specifically adapted to the cost- and performance requirements of the small satellite
market.
The new service, KSAT light, is initially composed of antennas ideally situated for serving polar orbits.
The locations at Svalbard (Arctic), Tromsø (Arctic) and TrollSat (Antarctica) ground stations, provide the
ability to communicate with, and download data from satellites practically two times per orbit (26 out of
28 possible daily passes). The polar sites will be complemented with low inclination sites as, where and
when user community demand develops. The new network is operated in parallel with the larger
diameter antenna network which remains indispensable to customers with higher requirements in
terms of download bitrates and absolute availability. The antennas used work in S-, X- and Ka-band and
have a diameter of 3.7 m.
At the same time, and closely associated with the idea of a high efficiency, low operational complexity
network, KSAT is rolling out an interactive web-based interface allowing the customer to instantly
visualize available pass opportunities across the entire network, book the passes and handle payments
and queries, all through a state of the art, intuitive graphic user interface.
KSAT light offers high quality and dependability and gives small satellite operators access to a reliable
ground station network without the need for initial investment as the pricing structure is entirely based
on a “pay as you use” approach. This minimizes financial risk for small satellite operators in case of
unexpected failures, short mission duration or premature end of mission. KSAT light makes using high
data rates for CubeSats easy and affordable.
57
Challenges and Solutions for the QB50 Telecommunication Network
G.March1, T. Scholz1, J. Thömel1, P. Rambaud 1 and S. Marcuccio2
1
2
von Karman Institute for Fluid Dynamics, Rhode-Saint-Genèse (Brussels), Belgium
Aerospace Division of the Department of Civil and Industrial Engineering, University of Pisa, Italy
QB50 space mission will provide the biggest CubeSat network in orbit. A constellation of 50 CubeSats in
a ‘string-of-pearls’ configuration will be launched together in January 2016 by a single rocket, into a
circular orbit at 380 km altitude. Due to the atmospheric drag the orbit will decay and progressively
lower layers of the atmosphere will be explored. Main goals are exploration of the lower thermosphere
with multi-point measurements, re-entry research and in-orbit science and technology demonstration.
In this analysis of communication functions the ground segment is analyzed, with a global overview of
different architectures, the main elements of a ground station, mission and control centres, and the link
between them.
This study is realized through the development of a tool which computes the number of stations
required to recover a certain amount of data generated by a constellation of satellites. This tool ensures
the efficiency of the communication system taking into account various design parameters like data
rates, limited elevation angles from ground stations, and the effects on the link quality such as orbit
perturbations, space and atmospheric losses and Doppler shifts. Particular attention is devoted to
frequencies: two different types of systems (UHF/VHF and S-band) are analyzed. In order to optimize the
positioning and number of stations, an iterative method is applied to compute the fraction of time when
a station is in view of a CubeSat in function of various parameters such as the latitude of the station, its
elevation and the altitude of the satellite. AGI-STK software was used to compute the access between
satellites in the constellation and ground stations, simulating system operability. Starting with a gradual
approach, the analysis begins by ideal study of the communication behaviour between one CubeSat and
a single ground station at VKI. After the selection of the communication system architecture, author
introduced the constellation concept, introducing also related considerations due to communication
overlaps between close satellites. Through this gradual investigation was possible to increase the
analysis complexity, having a detailed analytical description on communication behaviour, and finding
reliable results, which are extremely useful for mission accomplishment.
58
CubeSat Flight Experience
Lessons Learned
59
Lessons Learned from Four Months of UKube-1 in Orbit:
A Software Perspective
P. Mendham, M. McCrum
Bright Ascension Ltd, Scotland, UK
UKube-1 launched successfully on the 8th July 2014 and contact was made with the satellite shortly
after launch. At the time of writing this abstract the platform commissioning phase is in progress, and
there have been several interesting challenges to overcome during this period. As providers of the flight
and mission control software we have been involved with the UKube-1 project for more than two years
and a unique perspective on the mission. As the only team which spans both flight and ground
segments, from interfacing on board to satellite operations, we have learned many useful lessons. This
presentation reviews the mission development from a software perspective and presents a number of
key recommendations for other projects.
We begin by presenting an overview of the satellite development process, highlighting some of the
many challenges of UKube-1 and their implications on the system, and especially software, engineering.
These include the diverse payloads, demanding operational requirements and organisational challenges
from the large number of collaborating teams. We specifically focus on the spacecraft test process and
operations rehearsals, describing the evolution of the mission operational concept.
Next we describe the development of the ground segment architecture, centred around RAL Space at
Harwell in the UK. We describe the operational challenges and our software solutions, ranging from
automation for attended operations to a low-cost functional simulator.
Finally we describe the LEOPS and commissioning process and examine the reality of our various
software systems in use. Rather than going into technical detail, we describe each solution at a high
level, presenting the implications for the mission and recommendations for others.
60
The QB50 precursor flight: Status, preliminary results and lessons learned
Z. de Groot, J. Elstak, E. Bertels, J. Rottevee1
ISIS – Innovative Solutions in Space BV
On the 19th of June 2014, a Dnepr launch vehicle including the QB50 deployment system delivered the
two QB50 precursor satellites into their designated orbit. The two spacecraft, QB50p1 and QB50p2,
have been developed during an extremely compressed schedule. The on-ground development and test
phase was finalised within 6 months and the launch took place in under 9 months from the kick-off of
the project in October 2013.
The QB50 precursor team consists of ISIS – Innovative Solutions in Space, who is responsible for system
engineering, spacecraft design, Assembly, Integration and Verification (AIV) and the launch campaign
including the deployment system; Surrey Space Centre (SSC), who are responsible for the Attitude
Determination and Control System (ADCS); Mullard Space Science Laboratory (MSSL), who are
responsible for the Ion & Neutral Mass Spectrometer (INMS) Science Unit; Technical University Dresden
(TUD), who are responsible for the Oxygen Flux Probe (FIPEX); Ecole Polytechnique Federale de
Lausanne (EPFL), who are responsible for the Satellite Control Software (SCS); Astrium GmbH, who are
responsible for PA support; Von Karman Institute (VKI), who are responsible for the thermal analysis and
the Qarman thermocouple payload; Amsat Francophone and Amsat NL, who both delivered a radio
amateur communications payload.
The aim of the QB50 precursor mission is to de-risk key parts of the QB50 project and to learn valuable
lessons on the spacecraft and payload design, AIV and operations. This ensures that the participating
QB50 teams in the main mission can benefit from the lessons learned and experiences gained during the
precursor flight and thus minimizing risk in this challenging project. In addition, as QB50 will occupy the
radio amateur bands for a significant portion of the available spectrum during its mission (however
brief), the precursor satellites allowed AMSAT payloads to be embarked on the satellites as a ‘returnfavour’ to the radio amateur community for using their spectrum.
Since launch, the early operations and commissioning has focussed on getting both satellites in a stable
situation from a thermal, power, attitude and software point of view. Because of the extremely tight
development schedule, it was decided to develop the majority of the full mission flight software after
launch and subsequently using a software upload function to install the latest software images on the
spacecraft once finalized and tested thoroughly on ground. This meant that the software on-board at
the time of launch was limited to the most critical functionality. A lot of valuable information has been
acquired from both satellites, showing information about for instance the temperature environment
and power budget. Most progress has been made on commissioning the ADCS system and detumbling
the satellites to a stable Y-Thomson spin and commissioning of the payloads is currently about to start.
This presentation will explain the current status of the precursor flight, show preliminary results and
discuss the lessons learned for the QB50 main flight.
61
QB50 Precursor ADCS Flight Results
L. Visagie1, V.Lappas1, W.H. Steyn2
1
Surrey Space Centre, University of Surrey, Guildford, UK
2
Stellenbosch University, South Africa
The QB50 ADCS (Attitude Determination and Control System) will provide attitude sensing and control
capabilities to 2U CubeSats in order to meet the QB50 system requirements. CubeSats in the QB50
constellation require attitude control in order to minimize the influence of drag and to ensure science
payloads point towards the ram direction. The QB50 ADCS makes use of a combination of
magnetometer, sun and nadir sensor measurements and a MEMS rate sensor to estimate the current
attitude. It uses magnetorquers and a single reaction wheel, operating as a momentum wheel, to
stabilize and control the attitude of the satellite.
The QB50 Precursor mission aims to de-risk a number of technologies and programmatic aspects of the
main QB50 mission. The ADCS module is one of these technologies – each of the 2 precursor CubeSats
carry one of the QB50 ADCS modules, and operation of the precursor satellites involve commissioning of
the ADCS and stabilization of the platform by active control.
In the first few months, valuable experience has been gained in commissioning and operation of the
ADCS modules. In this presentation, the in-flight results of the precursor ADCS will be presented as well
as the problems that were experienced and lessons learnt.
62
A Worldwide Survey on the Regulatory and Economical Aspects of NanoSatellites
D. Andoni1, S. Cabrera1, K. Kornfeld1, P. Maier1, A. Raposo1, C.O. Asma2
1
Space Generation Advisory Council, Wien (Austria)
S-3, Swiss Space Systems, Payerne (Switzerland)
2
Development of nano-satellites is an emerging and key technological domain with more and more
nano-satellites being readied for launch every year. Other than their educational and PR benefits,
nano-satellites are also seen as complementary space technology platforms to bigger industrial
satellites.
This study, carried by Swiss Space Systems (S3) Department of Academic Projects and Space
Generation Advisory Council project groups, aims at understanding and showing the diversity of a
typical nano-satellite development, which would include technical, regulatory and economical
aspects. The project aims at performing a worldwide study of the typical roadmap and
considerations for the development of a nano satellite with an Earth-Observation payload. The
development concepts are to be analysed from the point of technical development, space law and
regulations, economics and business cases for the key technologies. Other than showing the diverse
options for handling all the developmental issues, this project also aims at proposing a feasible
roadmap for the future steps of the ‘small-scale’ organisations who would like to develop, test,
launch and operate their small satellites.
Answers will be sought for typical questions such as (but not limited to):





What is the typical cost of a nano-sat development, broken into items such as hardware,
human power, testing, launch, etc?
What are the typical regulatory steps that are followed by the nano-sat developers?
(existence of space law, registration, frequency allocation, definition of operator, definition
of launcher stage, launch contract, ITAR, import/export licences, insurances, etc)
What are the major differences among different countries/regions in terms of nano-sat
regulations?
How are the economic trends for development and the launch of nano-sats?
Would it be possible to propose an optimized strategy for the development, launch and
operations of the nano-satellites, taking into account the technical, regulatory and economic
factors?
63
LituanicaSAT-1: lessons learned from the first Lithuanian cubesat mission
L. Maciulis1, V. Buzas1, L. Bukauskas1, M. Dvareckas1, Z. Batisa1
1
Faculty of Mathematics and Informatics, Vilnius University, Lithuania
LituanicaSAT-1 is a 1U CubeSat developed by Vilnius University. The primary objective of the mission is
to provide university students and young researchers knowledge and real hands-on experience in
satellite design, development, deployment and operations. Hence the process helps to develop
infrastructure and know-how in space technology by interdisciplinary interaction between academia and
industry in Lithuania. Following key technology demonstrations have been carried out during the
mission:

Low resolution visible wavelength imaging system based on the open-source Arduino platform

Mode V/U FM transponder

In-house built silicon solar cells
Results from these technology demonstrations will be presented together with lessons learned from
mission operations and project management perspective. Special attention will be given to the feedback
received from amateur radio community that operated the FM transponder.
LituanicaSAT-1 is one of two Lithuanian CubeSats to be Lithuania's first satellites in space. The satellite
was launched from the International Space Station to space on February 28, 2014 by JAXA astronaut
Koichi Wakata.
64
Lessons Learned from Developing and Producing Structure and Mechanical
Systems for ESTCube-1
P. Liias1, E. Kulu2, M. Eerme1, P. Orusalu3, M. Noorma2
1
Tallinn University of Technology, Estonia
2
Tartu Observatory, Estonia
3
Protolab, Tartu Science Park, Estonia
ESTCube-1 is a 1U CubeSat which was launched on 7th May 2013 whose subsystems were developed by
students for educational purposes. It gave a unique experience and the educational outcome was much
better compared to using commercial products. The development and production of structures and
mechanical systems for nanosatellites generally receives little attention but there were many valuable
lessons learned in ESTCube-1 project.
ESTCube-1 frame is made from mono-block aluminium which was recommended by several persons and
publications early in the project to save weight and achieve required tolerances more easily. Considering
the production cost and complexity and the difficulty of systems engineering and satellite assembly it
might not have been the best choice. To meet the delivery date of the satellite last minute changes to
material selection had to be made. Some of the reasons were different legislative and material
numbering systems in Europe and USA, delivery dates and finances. What we have learned is that
problems caused by these kinds of changes may not come out during testing and can occur months after
a successful launch, for example screws will be magnetised and disturb attitude control. Due to mission
requirements a new antenna deployment system had to be developed for one side of the satellite which
needed to also have solar cells. Finding a supplier for beryllium copper antennas and burn wire for the
deployment was also challenging. For a long time ESTCube-1 had been expected to be launched with
PSLV from India when an opportunity to launch with European Vega rocket emerged. Vega has much
higher vibration levels during launch compared to PSLV which was unexpected. That put a lot of physical
strain on the harness which held the satellite during vibration testing which broke several times and
stronger ones were made after every test.
It is a long process from planning of the mission to launching a spacecraft, but it could be done much
more effectively when using the right methods, materials, manufacturing and testing. Learning from the
lessons of ESTCube-1 we have now started the development of new structure and mechanical systems
for the next missions of Estonian Student Satellite Program. Being part of the project since the beginning
has given a very good experience which components, materials and coatings should be used to achieve a
more streamlined development process.
65
Legal Aspects on CubeSats and Space Debris Issues
N. Antoni
Leiden University, International Institute of Air and Space Law
Small satellites are designed mainly for scientific or educational purposes at a low cost with a small size
and mass less than 500 kg. Many of them are based on the CubeSat standard and they are frequently
launched in low Earth orbit, with an increasing number of them being launched the last years -and
predicted to be launched in the near future- particularly into the Sun-synchronous orbit (SSO) at the
altitude between 650 and 800 km.
The major concerns that arise from these small satellite missions are the risks related to space debris
creation in SSO and in-orbit collision with other space objects. At the altitude of 650 and above natural
orbital decay decreases, and, as a result CubeSats remain in orbit longer than the life expectancy they
are designed for. Currently, their small size and mass prevent them to be equipped with a propulsion
system that would allow them to be deorbited and adhere to the 25-year in orbit lifetime limit as
defined by COSPAR and also required by their design specifications. In addition to this, experience has
shown so far a considerable failure rate of CubeSats launches either upon separation from the launcher
or after a sudden break-up or loss of contact only within a few weeks or even days of operation in orbit.
These non-functional small satellites are hazardous for the current and future inherently high risk
operations in the near-Earth space environment which is much crowded and congested, especially
after the collision of Cosmos 2251 and Iridium 33 satellites. The case of Cerise collision with an
untraceable object 20 years ago, when the population of debris was even smaller, should draw our
attention to this fact. Their inability to manoeuver and their very small size that might not be tracked
could pose an imminent risk not only to the mission itself but also to other space objects that cannot
exercise debris avoidance manoeuvres without having exact information of the location. The
aforementioned risks might be multiplied and, thus, intensify the problem in the case of many CubeSats
launched either in constellation or in formation flying in SSO, which will soon be the case.
In order to mitigate the above risks, it is highly recommendable for the operators of CubeSats missions to
comply with the voluntary UNCOPUOS Space Debris Mitigation Guidelines, which establish the legal basis
for a sustainable development of outer space activities. In accordance with this legal framework
CubeSats missions should be developed with the appropriate propulsion system that will allow them to
be maneuvered in case of collision risk and deorbited after the mission lifetime ends without exceeding
the 25-year rule. This will ensure the compliance of CubeSats operation with the corpus juris spatialis, in
the interest and benefit of the international space community.
66
E-st@r-I lessons learned and their application
G. Obiols-Rabasa1, S. Corpino1, R. Mozzillo1 and F. Nichele1
1
Politecnico di Torino, Torino, Italy
CubeSats are characterised to be small and cheap platforms, born within universities with educational
objectives. However, these systems are becoming more and more attractive for other missions, such as
for example technology demonstration, science application, and Earth observation. This requires an
increase of CubeSat performance and reliability, because educationally-driven missions have often
failed.
Nowadays, ESA Education Office is conducting its first edition of Fly Your Satellite! Program devoted to
provide support to selected University CubeSat developers of ESA specialists for verification phase of
their CubeSats. The goal of the initiative is to increase CubeSat mission reliability through several
actions: to improve design implementation, to define best practice for conducting the verification
process, and to make the CubeSat community aware of the importance of verification. Within this
initiative, CubeSat team at Politecnico di Torino developed the e-st@r-II CubeSat as follow-on of the est@r-I satellite, launched in 2012 on the VEGA Maiden Flight. Both 1U satellites are developed to give
hands-on experience to university students and to test an active attitude determination and control
system.
The present work describes the lessons learned gathered during e-st@r-I development and operations,
and their application to improve the new CubeSat, from design to operations. In particular, design
improvements have been applied to reduce assembly procedure complexity and to deal with possible
on-board computer failures. ECSS rules have been considered to design and assess new procedures for
the verification campaign, tailoring them when possible with the support of ESA specialists. Different
operative modes have been implemented to deal with some anomalies observed during the operations
of the first satellite; mainly leading to a new version of the on-board software. In particular, a new
activation sequence has been considered to have a stepwise switch-on of the satellite.
In conclusion, the know-how gained during e-st@r-I development and operations have been crucial for
the development and verification of the e-st@r-II CubeSat.
67
Orbital Dynamics, De-orbiting
And Debris Mitigation Techniques
68
The Aerospace Blockset for Xcos
P. Zagórski1
1
AGH University of Science and Technology, Kraków, Poland
Aerospace Blockset for Xcos is a free, open and extendable software tool for aerospace simulations. The
project was developed thanks to ESA Summer of Code in Space initiative. It is under active development
ever since 2012. From the very begging it was envisioned as a tool that fills the gap between the free
purpose-driven aerospace software (eg. Space Trajectory Analysis, Gpredict), and the very expensive
professional tools (eg. AGI STK, Matlab-Simulink). On the one hand anyone can download it and use it
for free for any purposes. On the other hand it is based on a powerful CelestLab library developed by
CNES and validated against true data and commercial simulation environments. It provides wide
functionality including but not limited to: propagating satellite orbits, propagating planet orbits,
conversions of reference and time frames, environmental models (Earth magnetic field, solar pressure,
atmospheric drag, etc...), ground station visibility, unit conversions, attitude dynamics and quaternion
algebra. It also provides real-time simulation capability to enhance ground station software suite.
Aerospace Blockset is a part of an open source Scilab/Xcos environment, which provides many other
compatible blocksets (FEM simulations, statistical analysis, etc.). It is also possible to freely modify and
write your own blocks to extend the capability and customize the tool for own needs. Users who
develop new functionalities are free to either keep them for themselves or contact the developers and
include them in future releases.
Its availability, capability, extendibility makes it a great tool for aerospace education and educational
CubeSat projects. Additionally, as the blockset is based on visual programming concept user can not only
perform complex simulations, but also better imagine and understand the nature of simulated
phenomenon by studying relationship between the functional blocks.
69
LitSat-1 Decay Analysis
V.Tomkus 1, D. Brucas2, D. Gailius3 , P. Kuzas3, A. Karpavicius3 and A. Vilkauskas3
1
2
Lithuanian Space Association, Vilnius, Lithuania
Space Science and Technology Institute (SSTI), Vilnius, Lithuania
3
Kaunas University of Technology (KTU), Kaunas, Lithuania
As CubeSat technologies get more and more popular due to its availability, low cost and short
development cycle, Lithuania has also joined the countries operating their own satellites in the
beginning of 2014 with two satellites LitSat-1 and LituanicaSat-1.
LitSat-1 satellite has been developed as the test and demonstration platform for future mission, to
analyse the possibilities and ways of constructing, launching and controlling CubeSats. Due to that
the construction design was as simple as possible. Considering the further research of Attitude
Determination and Control (parameters important for that), 1U satellite had a mass of 950 g, passive
magnetic attitude control (permanent magnets and Permalloy), and three 10 mm wide
communication antennas in VHF and UHF bands.
The satellite was launched on 28th February 2014 from ISS (at an altitude of 420 km) together with 4
other CubeSats, and became one of the two first orbital bodies manufactured in Lithuania. After the
extremely successful orbital experimental work, the satellite faced an extremely rapid (both in time
and compared to other satellites launched at the same time) orbital decay with orbital life of only 83
days. Such rapid decay could be contributed to low ballistic coefficient of satellite (high drag and low
mass), nonetheless it could not be explained only by the influence of aerodynamic drag. According
to comparison of theoretical (computational) and experimental data it was determined that besides
aerodynamic drag, the rapid decay of the satellite was caused also by passive magnetic hysteresis
damping of the Permalloy having quite considerable influence on the dissipation of the kinetic
energy .
Due to relatively long two 145 MHz antennas the 1U Cubesat attitude was determined by the
prevailing sum of aerodynamic and gravitational torque vectors combined with changing direction of
the magnetic vector. It caused the complex behaviour and spinning of the satellite around the mass
centre with average period of 18 s and remagneting of the Permalloy (mass of 6 g). According to the
theoretical calculations such remagneting decreased orbital speed (and caused decay of satellite) by
average Δv of 5x10-3 m/s per hour. These assumptions together with further analysis of data and
comparison with other satellites (launched at the same time) will be given in the presentation.
70
In orbit testing of a de-orbiting sail on the Cubesat URSA MAIOR
M. Valdatta1, N. Bellini1, A. Locarini1, S. Naldi1, D. Rastelli1, F. Piergentili2, F. Santoni2
1
2
NPC S.r.l., New Production Concept/Spacemind division, Italy
CRAS, Centro Ricerca Aerospaziale la Sapienza, University of Rome “La Sapienza” Italy
One of the most important innovations in space sciences, due to the recent increasing of cubesat
based missions, is the possibility to have a low cost platform suitable for in orbit testing of new
technologies.
Many universities continue to build cubesats for educational purposes so that the number of these
nanosatellites launched into space is rapidly increasing. Cubesats are, in general, not provided
with a post mission disposal system to perform de-orbiting at the end of operative life. The result
is an increasing number of space debris in a size range which is at the same time difficult to track
and potentially destructive for operative missions. Moreover, typical cubesat features do not
include active attitude and orbital control systems capable to perform a de-orbiting manoeuvre.
The space on board of these nanosatellites is a particular issue especially for 1U cubesats, where
usually all the available space is filled with the main subsystems of the satellite itself. Therefore,
post mission disposal systems do not play a key role for operative life of cubesats and are not
considered as design drivers. In case of existing de-orbiting guidelines becoming rules, a post
mission disposal system would become mandatory for all satellites, including nanosatellites. For
this reason, the Spacemind division of NPC Italy has designed and manufactured a de-orbiting drag
sail which will be host onboard the 3U cubesat Ursa Maior part of the constellation of QB50. The
sail is a square of 70x70cm based on the use of a polyurethane foam that can be stored in a little
volume. Once the closing mechanism is released the sail returns to the original shape. The idea is
to use the mission for in-orbit validation of the system. The post mission disposal device
constitutes a stand-alone system with its own electronics and power resources, nonetheless a
connection with the satellite power supply has been considered as a backup in case of failure. In
general the sail is designed to guarantee the respect of the IADC guidelines for all the orbits
commonly used in a nanosatellite mission.
71
A tether-based aerodynamic de-orbiting system
O. Vallet1, C.O. Asma2
1
2
ELISA, ÉcoLe d’Ingénierie des Sciences Aérospatiales, Saint Quentin (France)
Swiss Space Systems, Payerne (Switzerland)
Keeping the near-Earth orbit in a sustainable condition is the main purpose of the 25 years lifetime
rule for LEO (Low Earth Orbit) satellites. CubeSats orbiting under 600 km naturally respect the
criteria but the others have to integrate a post-mission disposal system. The main issue is to deorbit
the satellite with an efficient passive system, which will use minimal power, mass and volume
resources. In this way, a “tethered sail” system is under study to efficiently allow the deorbiting of
LEO CubeSats.
The studied concept involves tethers or tether-like sails with a small mass at their tips, simply
deployed from the satellite at the end of their functional lifetimes. With no rigid structures to
maintain the drag area but by making use of the natural gravity gradient positioning of tethers, the
deployment issues of common solar sails are reduced. The length of the system will only be metric
scaled (not kilometric like usual tethers) which will lead to reduced deployment, orbital debris and
oscillations issues.
There is no use of any Electromagnetic or Electrodynamic effect, but only the drag capabilities. The
system will be useful until the drag forces clearly prevail against the gravity forces. For a 3U CubeSat,
the objective could be to reach 550km. From this point, when tumbling, the satellite re-enters within
4 years even in case of system failure.
The improved area-to-mass ratio for such a system when applied on CubeSat makes it potentially
interesting. The endpoints masses of the tether can be some of the satellite’s original subsystems
(solar panels, structure…), in order to not launch useless mass in orbit.
Presentation will approach the work under development of potential mechanisms, material, and
results of the deorbiting system for different typical missions.
72
De-risking Active Debris Removal with
CubeSat in-Orbit Demonstrations
M. Richard, G. Feusier, R. Wiesendanger, C. Pirat, C. Paccolat, F. Belloni, D. Courtney*
A. Pollini**
D. Bovey, J. Buchli, M. Bircher***
* Swiss Space Center, EPFL, Switzerland
** HES-SO, Switzerland
*** ETH-Zurich, Switzerland
This presentation will be discussing the current results of the analysis performed in the frame
of the ESA study « CubeSat Technology Pre-Developments, QB-50- Active Debris Removal”.
Two mission scenarios utilizing CubeSat technologies for the main satellite subsystems have
been investigated. These two missions are called CADRE, for CubeSat Active Debris Removal
Experiment. The first CADRE mission scenario involves the demonstration of rendezvous
sensor technologies and operations. The second CADRE mission scenario is targeted toward
the demonstration of net deployments and flexible link (tether dynamics). Both CADRE
missions assume the use of 6U (6 Units) to 8U chasers and targets of 3-4U, launched together
in a 12U deployer. The analyses performed include a trade-off of the vision-based rendezvous
sensors and of the radar technologies, net technologies, and also trade-offs at the CubeSat
subsystem levels, especially for micro-propulsion. The outcome of these analyses includes
conceptual mission and satellite designs, technology gaps identification and make-or-buy
options, and a set of mission, system and technology requirements.
One of the key aspects of the CADRE IOD is the scalability and applicability of the
demonstration utilizing such small flight systems. These aspects are addressed for the
rendezvous sensors (in the three primary functions of Debris Detection, Debris Identification
and Debris Motion Reconstruction), for the Guidance, Navigation and Control, for the Net
capture system and its dynamics, for the flexible link tether dynamics, and for the mission
operations. Considerations and preliminary scalability conclusions will be presented.
Current issues and limitations with such IODs will be discussed, as well as conclusions on the
feasibility of CADRE missions.
73
CubeSat Networks and Constellations,
Formation Flying
74
Space-based Ad hoc network:
a solution for multiple satellite TT&C problem in QB50 project
Pengfei Liu1, Lei Yang1and Xiaoqian Chen1
1
Institute of Space Technology, National University of Defense Technology, China
For the QB50 project, 50 CubeSats will be launched and injected in 2016 into a 380km orbit to perform
in situ observations of the thermosphere as well as science and technology in-orbit demonstration.
Since most of these satellites have no position control, their orbit trajectories are mainly subject to pure
orbital mechanics. Simply relying on the slight differences in orbital elements caused by in-orbit release
mechanism, these satellites are not easy to drift apart in short periods after deployment. That is to say,
more than one satellite would be located in the boresight range of one ground station antenna at the
same time. Since some of the QB50 satellites may adopt the same up- and downlink frequency, the
TT&C link would be affected due to co-channel interferences. How to keep an efficient TT&C channel
between all the 50 CubeSats and ground stations during initial orbital periods is a difficult problem for
the system design of this project, and is also the topic of this contribution.
The idea to solve the multiple satellite TT&C problem in QB50 project by space-based ad hoc network is
proposed in this contribution. An integrated system architecture of space-based ad hoc network is
described. To support the rapid initial networking process of satellite cluster when just deployed by the
rocket, a GPS aided network formation algorithm is presented. On the base of IEEE802.15.4 standard, a
light-weight network protocol stack is designed with a cluster-tree routing algorithm and a smart
network address distribution mechanism. Result from the simulation scenario built in OPNET has
demonstrated the efficiency of our design.
A system-in-loop platform is being built for further perfection of the system design. This paper describes
technical details, hard- and software design, and also addresses some related results of this platform.
75
TW-1: A CubeSat constellation for space networking experiments
Shufan Wu*，Zhongcheng Mu*，Wen Chen*，Pedro Rodrigues**, Ricardo Mendes**
*Shanghai Engineering Centre for Microsatellites, Chinese Academy of Science, Shanghai,
China
**TEKEVER Space, Portugal
In the past decade, Cubesat has gained more and more attentions in space communities, has
evolved from purely educational tools to a useful platform for technology demonstration and
scientific instrumentation, and has walked out of university labs into many potential
applications. Networking and constellation with multiple Nanosats and CubeSats are
foreseen an important direction for different applications. On this topic, two state of the art
communication technologies, the software defined radio (SDR) based inter satellite
communication and the ad hoc adaptive networking technologies, enter into the front stage.
This paper presents a small space networking experiment mission (TW-1 project), to test and
validate the new devices, and explore the cubesat space application based on the two
technologies above discussed. The major technologies and related instrument or device
modules to be used in this mission are GAMALINK, which is an S-band inter-satellite
communication module, a novel dual band GPS/BD receiver, an AIS receiver, and an ADS-B
receiver, all being designed based on SDR technologies. Also a novel cold-gas micro
propulsion module based on MEMS technology will be used for orbit and constellation
control.
TW-1 project consists of three CubeSats carrying different payloads and instruments with
one 3U CubeSat and two 2U CubeSats, to be put into an LEO orbit, forming an along-trace
satellite network and/or constellation. It is designed and being implemented by a consortium
led by the Shanghai Engineering Centre for Microsatellite in China, together with GomSpace
from Denmark, Tekever Space from Portugal and NanoSpace from Sweden. The main tasks of
this mission are listed in the following:
 Objective 1: CubeSats networking based on Gamalink；
 Objective 2: Monitoring sea ice and gaining the maritime traffic information in polar
region based on AIS receiver and camera；
 Objective 3: Demonstration of autonomous formation flying including the
along-track orbital (ATO) formation and the projected circular orbital (PCO)
formation；
 Objective 4: In-orbit demonstration and validation of ADS-B receiver/ Gamalink /
Micro-propulsion；
 Objective 5: Imaging the satellite separating process.
76
Status of the QB50 Project
J. Thoemel, F. Singarayar, T. Scholz, C. Asma, P. Testani, D. Masutti, J. Muylaert,
von Karman Institute for Fluid Dynamics, Rhode-Saint-Genèse, Belgium
CubeSats have emerged to be recognized powerful tools for a new class of space missions. They have
served many objectives and mostly to educate young space engineers by means of the hands-on design
and manufacturing experience.
The QB50 project aims at the use of the CubeSat concept to further facilitate access to space for the
future generations, to conduct unprecedented science, to demonstrate new space technologies and also
to provide training to young engineers.
To this end, the Project, coordinated by the von Karman Institute for Fluid Dynamics, Belgium, has
invited universities from all over the world to submit a proposal for a CubeSat to be embarked on the
mission.
The QB50 consortium is managing the mission and in particular it develops the deployment system, the
common sensors that will be placed on all science satellites, and procures the launch service. In
addition, it provides a number of key technologies and services such an attitude control system and a
satellite control software.
A number of such are being tested on the QB50 precursor mission.
Started in November 2011, the project is now beyond the detailed design phase. All technologies
developed by the consortium and community have appeared now at as hardware for display,
demonstration, test or even flight purposes. The project now prepares for the assembly of the satellites
and the deployment system.
The QB50 consortium consists of the following partners: ISIS B.V. (NL), MSSL/UCL (UK), EPFL (CH), SSC
(UK), B.USOC (B), TU-Delft (NL), IAP (D), DLR (D), Stanford (US), ITAM (Russia), NPU (China), Airbus
(D/Fr), SSLLC (USA), VKI (B). Sensor Unit providers are UiO (Norway) and TU-Dresden (D). The project is
partially funded by the European Commission Framework Program 7 Grant 284427 and by consortium
and community CubeSat provided in-kind financial contributions. It has established Memorandum of
Understandings with AMSAT (UK/NL/Fr), Aalborg University (Denmark) and the SGAC and received
substantial support from governmental organizations such as BELSPO (B), BIPT (B) and the ITU. Most
importantly the project consists of 50+ CubeSat developing teams. The highly motivated work of
individuals and organizations is very much appreciated by the authors.
77
Attitude Determination
and Control
78
Aalto-1 Nanosatellite Attitude Determination and Control System End-to-End
Testing
T. Tikka1, O. Khurshid1, N. Jovanovic1, H. Leppinen1, A. Kestilä1 and J. Praks1
1
Aalto University, School of Electrical Engineering, Helsinki, Finland
In this paper we present a hardware-in-loop (HIL) test setup and usage designed for high performance
Attitude Determination and Control System (ADCS) end-to end testing and validation for multi-payload
Cubesat missions.
The Aalto-1 mission requires accurate pointing of a miniature radiation monitor and a hyperspectral
imager, and a 200 °/s spin-stabilized operation mode for an electric solar wind sail based plasma brake
experiment.
The satellite’s ADCS, iADCS-100 provided by Berlin Space Technologies (BST), contains sensors and
actuators typically only seen in larger satellites: star tracker, gyroscopes, a magnetometer,
magnetorquers and reaction wheels. In addition, six digital sun sensors and a GPS receiver are
integrated to the system.
To verify correct operation of the ADCS before launch, and to assure compatibility with the satellite’s
scientific mission, a thorough testing campaign is currently being performed. BST conducts
environmental qualification, functional testing and control algorithm testing, whereas end-to-end
mission tests and acceptance tests are performed in Aalto University.
The mission tests are carried out using a HIL test setup running an attitude and orbit dynamics simulator
in Simulink xPC Target. The simulation provides real-time sensor data to the ADCS through an I2C
interface according to simulated sensor and actuator models, mission operations and disturbances. By
connecting the HIL setup to the ADCS while integrated to the rest of the satellite subsystems, even the
most complex mission operations can be tested and validated end-to-end in a closely flight
representative configuration.
79
Star Tracker Cost Reduction for Small Satellites
Tjorven Delabie1, Joris De Schutter1, and Bart Vandenbussche2
1
Department of Mechanics, KU Leuven, BE
Institute for Astronomy, KU Leuven, BE
2
In recent years, the great potential of small satellites has become ever clearer and small satellites
are selected to perform increasingly complex missions. With this rise in mission complexity, the
requirements on the Attitude Determination and Control System of the satellite increase as well.
Of all the attitude determination sensors, the star tracker is by far the most accurate one. The
accuracy of this sensor is in the order of arc seconds. The disadvantages of this sensor are that it is
expensive, takes a considerable volume, and has a high power consumption. In this paper, we will
discuss the star tracker developments that are currently being done at the KU Leuven University.
These star tracker developments are part of the development of an ADCS for the SIMBA Mission,
which is scheduled to launch within the QB50 campaign.
In the first part of this paper we discuss how the novel star tracker algorithms developed at KUL can
reduce the cost of the Star Tracker. Both the centroiding algorithm and the tracking algorithm have
a significantly reduced computational cost, thanks to analytical solutions of the optimization
problem. This can allow to save costs in the electronical hardware and will reduce the strain on the
power budget. Furthermore, the star identification algorithm and tracking algorithm are
significantly more robust to inaccurate measurements. This allows to yield high accuracy, even with
lower cost components. The algorithms will be presented and we will focus on the increased
efficiency.In a second part, we discuss the tests that are performed to analyse the performance of
the star tracker. For small satellites, testing procedures are often not as standardized as they
generally are for satellite missions. As the SIMBA CubeSat is currently being developed as ESA’s first
CubeSat through an ESA GSTP project, the test campaign of the KUL star tracker will adhere as
strictly as possible to the standards set by ESA. The procedures that are followed will be outlined in
this paper and may serve as a guideline for future star tracker test campaigns. This may help to
reduce the time and money needed to devise and set up a test campaign for future missions. Since
setting up a test campaign is often a serious strain on the manpower and financial budget, this
could lead to a serious reduction in cost and lead time. An outlined procedure would also facilitate
the comparison between different star trackers on the market and would allow small satellite
developers to select the best star tracker for their mission.
Both the novel star tracker algorithms and developed testing procedures will allow to make the
accurate star tracker more accessible for small satellites. The increased attitude knowledge accuracy
that this sensor brings to the satellite platform will allow small satellites to perform even more
complex and interesting missions. This will again lead to new opportunities and new developments
for this growing group of satellites.
80
ZA-AeroSat: A QB50 CubeSat demonstrator for multidisciplinary technology and
scientific research
M. Kearney1, W.H. Steyn1
1
Electronics Systems Laboratory, Stellenbosch University, Stellenbosch, South Africa
CubeSats have proven to be valuable tools that can be used to satisfy a wide variety of commercial and
scientific demands. ZA-AeroSat is designed with both of these goals in mind. The satellite is built by a
team at the Electronic Systems Laboratory (ESL) at Stellenbosch University. It will be launched into a
<400km Lower Earth Orbit (LEO) as part of the QB50 CubeSat constellation; an international project
which is led by the Von Karman Institute for Fluid Dynamics (VKI).
In terms of technology demonstration, the satellite will carry a variety of local South African CubeSat
subsystems. The Attitude Determination and Control System (ADCS) of the satellite is the main
subsystem designed by the ESL. This highly integrated, compact ADCS unit will be flown on-board more
than 10 other QB50 satellites as their main ADCS. The first two units are flying currently on-board the
QB50 precursor satellites, which serves as testing platforms for the technology being flown on QB50
satellites. The ESL developed star tracker, CubeStar, will be flown as payload and possibly used as part of
the ADCS system. The satellite will also feature ESL designed mechanics in a deployable magnetometer
and deployable 2U solar panels. These solar panels will be deployed to a specific angle, chosen to
provide optimal power to the satellite throughout the entire range of Local Time of Ascending Node
(LTAN) values the satellite will experience. This is a particular challenge considering the significant LTAN
variation due to its fast decaying LEO orbit. Communications hardware built by a local partner, Cape
Peninsula University of Technology (CPUT), will also have its maiden flight on board ZA-AeroSat.
ZA-AeroSat will carry 3 scientific payloads. The first is the QB50 FIPEX science unit. The sensor will
measure atomic oxygen in the lower thermosphere. This will, among other uses, serve as validation for
current atmospheric models. Further, it will monitor the temperature of the different surfaces of the
CubeSat to compliment the science data collected. The second scientific payload is the satellite’s
featherlike antennas. The 4 VHF/UHF antennas will act as passive aerodynamic stabilization elements,
while simultaneously enabling a communication link between the satellite and the ground. The satellite
will use its ADCS unit in combination with these antenna feathers to control its orientation to within 5 ͦof
its orbital velocity direction throughout the mission lifetime. Lastly, the satellite will carry a scientific
experiment which aims to utilize a novel method to measure gravitational waves. This device will be
able to measure the gravitational effects of the earth, sun and moon.
81
Attitude Control Simulation Using Variable Speed CMG for 3U CubeSat
H. Kim1, H. Lee1, and Y. Chang2
1
Space System Research Laboratory (SSRL), Korea Aerospace University (Goyang), Republic of Korea
2
School of Aerospace Engineering, Korea Aerospace University (Goyang), Republic of Korea
Control Moment Gyros (CMG) help small satellites become more agile and manoeuvrable. CMG has
been implemented for CubeSat recently as the hardware miniaturization is feasible and its use in small
satellite is being steadily increased. KAUSAT-5, a 3U CubeSat under development by Korea Aerospace
University, will implement an advanced Variable Speed Control Moment Gyro (VSCMG) to demonstrate
its capabilities in orbit. VSCMG is able to generate torques by changing flywheel speed with fixed gimbal.
This research is regarding attitude control simulation according to each of the operation modes of
KAUSAT-5 with a mounted VSCMG. One of KAUSAT-5’s missions is to demonstrate VSCMG in a space
environment for future applications. Gimbal angles need to be determined using relative encoder in the
early phase of the satellite, because they can become biased by vibrations in a launch environment.
After that, gimbal angles should be reoriented using null motion for their regular operation. In mission
mode, attitude control for targeting a desired point on an uploaded schedule is performed. In this study,
the attitude control modes for each operation mode of KAUSAT-5 are defined, and equations of motion
are derived with nonlinear techniques by considering structural properties, CMG specifications, and the
CMG cluster configuration of KAUSAT-5; adaptive control laws at each operation modes are also
suggested. Attitude control simulation is performed based on the derived equations of motion and
control laws under MATLAB/Simulink. The paper also shows through numerical simulation that the
designed control laws are applicable to ultra-small satellites.
82
The piNAV-L1 – The World’s First Ultra Low Power CubeSat GNSS Receiver
J. Laifr1
1
SkyFox Labs, Prague, Czech Republic
The piNAV-L1 is the World’s First Low Power Space-Friendly CubeSat GNSS receiver specially
designed to provide continuous accurate position determination in LEO onboard small satellites or
high altitude balloon missions with limited power and mass budgets. It requires only a fraction of
power
(150 mW maximum) in comparison with conventional space-grade GPS receivers allowing
permanent
1 Hz data output within typical 1-Unit CubeSats.
Easy-to-use serial (UART) data interface output providing standardized NMEA messages together
with external GPS antenna provides a smart standalone solution for all kind of space-grade projects
where the precise position, time, date and velocity information is needed. The PPS (Pulse-perSecond) and PF (Position Fix) signals are available on System Interface connector to indicate the
receiver status.
Ultra low mass and dimensions (753511 mm) fits perfectly with all kind of space-demanding
satellite projects. Additional aluminium radiation shielding is delivered with the Space-Grade Flight
Models (FM). High altitude, high velocity Engineering Model (EM) with identical mechanical and
electrical properties is available with software limitation at reduced pricing.
The receiver has been successfully tested onboard the stratospheric balloon flight up to 35 km of
altitude and the space flight test is scheduled for the winter 2015.
83
In-House Magnetic Field Simulator For Cubesats
M. Balan1, C. Dragasanu1, M. Pripasu1, S. Radu1, C. Cherciu1, M. Trusculescu1
1
Institute of Space Science, Magurele (Bucharest), Romania
The presentation presents a three axis controlled Helmholtz cage designed as an attitude
determination and control test bed for CubeSat type nanosatellites. Having a useful testing volume
of approximately 30 dm3 and a maximum designed magnetic field intensity of 150 μT for each pair of
coils, the cage can simulate the entire range of LEO orbits.
Starting from the coil design mathematics, the paper presents in detail the magnetic field measuring
and control equipment chain. Initially, the magnetic field intensities values along the orbit are
determined by using STK software and translated in real time to current intensities on each pair of
coils. Moreover, a custom SGP4 orbit propagator is implemented and the magnetic field intensities
value can be obtained by IGRF interrogation.
The system can work in most common configuration where the magnetic field is generated in an
Earth inertial reference frame or Earth Centered Earth fixed reference frame simulating in this way
the magnetic field vector. Moreover, a detumbling test mode has been implemented. In this mode,
the satellite rotational matrix is added to the satellite trajectory in the STK simulation scenario and
the magnetic intensities on each axis are generated in the body reference frame. The magnetic field
computed in this way is generated with the cage having the satellite fixed inside. This rapid variation
field is used to trick the satellite magnetometer and observe the satellite behavior in the detumbling
operational mode.
The cage is powered from a current source with a resolution of 0.1 mA which makes possible the
obtaining of 10 nT of magnetic field increments on each axis. The cage control can be done in an
open loop configuration by supplying a proportional current or with a tri-axis magnetometer used as
sensor for the control loop. The paper concludes with the experimental data obtained during the
cage magnetic environment characterization and the zero offsets determinations.
84
Active magnetic attitude control algorithms for CXBN-2 CubeSat
M. Ovchinnikov1, V. Penkov1, B. Malphrus2, K. Brown2, D. Roldugin1
1
Keldysh Institute of Applied Mathematics of RAS, Moscow, Russia
2
Morehead State University, Morehead, KY, USA
Magnetic attitude control system for CXBN-2 satellite is considered. The goal of the CXBN-2 mission
(follow-up of successful CXBN) is to increase the precision of measurements of the Cosmic X-Ray
Background in the 30-50 keV range to a precision of <5%. CXBN-2 has already been selected for flight
through the NASA ElaNa Program and will be launched in 2016-2017.
Control system should provide possibly maximum number of data sets for the payload and possibly even
celestial sphere coverage. Absolute minimum data volume is million seconds of scientific data per year.
The main source of data loss is sensor sensitivity to high-energy illumination from close sources. Earth,
Moon and Sun blanket exposure in case located in sensor field of view. The sensor itself is resistant to
this illumination and is not damaged, however the data are lost. Four different control strategies are
proposed and studied. They are assessed according to a number of criteria, ranging from the scientific
mission requirements to engineering and mathematical robustness of a system. Following modes are
studied:
1.
Spin-stabilized satellite with regular spin axis rotation. The satellite is considered axisymmetrical
one, payload sensors are perpendicular to the spin axis. This provides continuous rotation for sensors
field of view.
2.
Spin-stabilization with Earth avoiding. Spin axis is always directed roughly to the local vertical.
3.
Spin-stabilized with solar panels charging and Earth avoiding. Satellite faces the Sun for half an
orbit to provide battery charge. Then it moves to local vertical stabilization.
4.
Free-flying satellite with speed control. This regime provides no specific attitude. Angular
velocity is affected only, the satellite should keep rotating to cover full sphere.
Free flying with occasional (de)tumbling is proved to be the best solution in comparison with different
spin stabilization schemes. Although the simulations modeling the free flying concept of operations will
result in less effective sky coverage near the polar regions, this is not anticipated to be a problem given
that the science data is expected to be less usable in these regions. The expected scientific data gain is
present.
85
Characterisation of Hysteretic Dampers for
Passive Attitude Control of CubeSats
D. Ivanov1, V. Penkov1, D.Roldugin1, M. F. Barschke2, K. Briess2, N. Kupriyanova1
1
Keldysh Institute of Applied Mathematics of RAS (Moscow), Russia
2
Technische Universität Berlin, Germany
As many CubeSat mission scenarios do not require active control of the satellite’s attitude, passive attitude control is considered an efficient solution due to its simplicity. Such systems typically consist of a
strong permanent magnet and hysteretic damper made by a special soft magnetic material with low coercivity and high magnetic permeability at low external field intensity that reaches saturation in the geomagnetic field. The magnet aligns the satellite to Earth’s magnetic field lines, whereas the damper is
required to reduce the oscillation. Magnet and damper must be carefully matched in order to ensure
proper functioning of the system. While dimensioning of the permanent magnet is comparatively simple, the damping capabilities of the hysteretic damper are not only dependent on the amount of material used, but also on the shape of the damper. Therefore experiments are required to determine the
actual damping properties for a certain design.
Technische Universität Berlin developed a hysteretic plate damper for passive satellite attitude control
within the BEESAT CubeSat series. While a plate shaped damper was expected to be less preformat than
a rod system, it is significantly easier to integrate, since it has the same form factor as the satellite’s
electronic boards. Within the framework of a bilateral agreement, characterization of prototypes of this
damper was conducted at the Keldysh Institute of Applied Mathematics.
This paper presents the results of theoretical study and laboratory experiments on hysteresis plates and
rods that were conducted to obtain and compare their damping capabilities. Coercivity, permeability
and saturation remanence of various damper designs were derived from an experimentally determined
hysteresis loop curve of the damper. These properties were then used to estimate the ohmic and hysteresis energy dissipation at different angular velocities. In order to evaluate and compare the effectiveness of the different damper designs, the obtained properties were used for numerical simulation of
CubeSats equipped with different dampers for various angular velocities. Hereby, a comprehensive collection of suitable design, carefully adjusted to different use cases is presented.
86
A Constrained Attitude Control Method for Aoxiang-Sat
R.Liu, H.Chen, Y.Liu, X.Yu and J.Zhou
Shaanxi Engineering Laboratory for Microsatellites,
Northwestern Ploytechnical University, Xi’an, China
Aoxiang-Sat is a 12U CubeSat aimed at performing scientific and technological experiments in-orbit
and scheduled to launch in 2015. It is an interdisciplinary, interdepartmental effort that has
involved over 30 teachers and students, researched by Shaanxi Engineering Laboratory for
Microsatellites of NPU. The main goals of the Aoxiang-Sat are detecting the skylight polarization
patterns and measuring gravity. These main tasks request the satellite to be earth-pointing and
three-axis stabilization.
The Attitude Determination and Control Subsystem (ADCS) is the precondition of fulfilling all kinds
of tasks. Many commercial products are selected to complete attitude tasks. In order to meet the
targets of being cheaper and lighter, the numbers of sensors and actuators are as few as possible.
The sun sensor we used has a field of view (FOV) of 60 degree. Orbit analysis indicates that at least
three sun sensors are needed to achieve the attitude demands. Finally, three sun sensors, three
gyros and three magnetometers are selected to accomplish attitude determination, three flywheels
and three magnetorquers are used as actuators. The ADCS of Aoxiang-Sat can realize functions such
as earth-pointing, sun-pointing and three-axis stabilization.
Since the Aoxiang-Sat just has three sun sensors, there are many cases that the sun can not be seen.
This paper presents an attitude control strategy for three-axis stabilization CubeSat with
constrained FOV of sun sensors. For the sake of stabilization, sun acquisition must be done first,
which is the first difference from common methods. Attitude determination begins after sun
sensors capture the sun. The results of attitude estimation are used to undertake attitude
stabilization. An attitude path is programmed to ensure that the sun sensors can track the sun all
along the maneuver, which is the second distinction from others. Simple but effective control law is
applied to decrease the load of On-Board Computer and to guarantee high reliability. Numerical
simulations demonstrate the effectiveness of the proposed attitude control strategy.
87
Future Technologies
on CubeSats
88
Characterization & Design of Solid State Hinges for
Deployable Cubesat Structures
Elbara Ziade* (eziade@bu.edu), Calvin Patmont* (cpatmont@bu.edu),
Nathan Darling (thomas77@bu.edu), Theodore Fritz* (fritz@bu.edu)
Center for Space Physics, Boston University, Boston MA USA
*
Authors contributed equally to this work
The success of the university-developed CubeSat specification has already demonstrated the broad
impact that a simple, robust and modular satellite bus has on military, civilian and private industry
spaceflight endeavors. In addition, the economic benefits of configuring and qualifying commercial
off-the-shelf (COTS) technology for space applications continues to increase the popularity of
Cubesat missions. However, as Cubesats evolve their electronic and instrument complexity to
achieve more sophisticated endeavors, the constrained power budget has to be met by using either
more expensive solar cells that are outside a Cubesat budget, or by incorporating deployable solar
panels to increase the sun-facing surface area. Often small satellite teams design an in-house solar
panel deployment system because available space-rated deployment systems are too expensive.
Therefore, there is a need to design an economical deployment system that is within a Cubesat
financial constraints.
Our response is a scalable and configurable hinge for popular Cubesat sized satellites. In lieu of the
common pinned-joint torsional-spring hinge - which must account for the effects of thermal
expansion, galling, cold welding and lubrication in a vacuum environment - we have developed a
“solid-state”, or a non-linear single-component hinge. Through a novel off-planar arrangement of
COTS spring steel strips, we have designed a self-actuating, self-guiding, and self-locking hinge; while
circumventing the risks associated with the pinned-joint scheme.
In this paper we investigate the dynamic response of the spring steel hinge to find: 1) the momentangle relationship the hinge places on common solar panel form-factors; 2) the impulse that the
locking solar panel places on the satellite structure; 3) and the damping coefficient internal to the
hinge. These are obtained through a combination of tests on typical 1U, 3U and 6U Cubesat solar
panel form-factors. We performed benchtop static tests that measure the static-moment vs
deployment angle; and dynamic zero-gravity tests on a NASA reduced gravity flight that measure the
dynamic moment vs. deployment angle and dampening time. Using the results, a satellite engineer
is able to configure a spring steel hinge for his/her solar panel’s form factor and attitude
determination capabilities, without extensive testing requirements.
89
Autonomous Command and Data Handling System for a 3U CubeSat
L. Feruglio1, R. Mozzillo1, S.Corpino1 and F. Stesina1
1
Politecnico di Torino, IT
Over the last few years, increasing efforts have been spent by the scientific community on enhancing
the autonomy of a space mission, both concerning ground and space segment. Different solutions have
been proposed, from satellite procedure execution languages, which aim to reduce the chances of
operator mistakes and improve monitoring, to the implementation of on-board autonomous
capabilities, for health keeping, efficient resource optimization, communication planning, and more.
The paper presents the Command and Data Handling (C&DH) subsystem of 3-STAR, a 3U CubeSat
project currently under development at Politecnico di Torino. The program started to take part in the
GEOID constellation promoted by the European Space Agency, but it later evolved to be used also as a
stand-alone technology innovation platform: this is the reason why the satellite is being developed
keeping in mind future adaptation to different types of payload.
In this sense, the 3-STAR C&DH design features algorithms for autonomous decision making and healthkeeping, based on functions of pattern recognition and intelligent machine learning.
A mission case study for 3-STAR is presented, showing a comparison between traditional mission and
envisaged autonomous operations, highlighting critical aspects of both technologies and detailing the
improvements gained through the use of artificial intelligence on-board. In addition, the architecture of
the C&DH, described using model-based representation, is depicted for both the on-board hardware
and the algorithm itself.
Results show how boosting on-board autonomy can greatly improve mission reliability especially for
CubeSats, where a continuous communication link cannot always be granted, and where the teams
involved cannot usually allocate many human resources to the ground control station. In addition, onboard autonomy allows to satisfy mission operation design criteria that would be significantly more
demanding using traditional approaches: fast response to critical events, relevant telemetry downlink,
on board scheduling optimization, are among the features that will be within reach of an increasing
number of teams, especially, but not limited to, university ones.
90
Early Orbit Phase of Deployment Mission of Inflatable Membrane Structure of
Nano-Satellite ''SPROUT''
K. Mita, M. Yamazaki and Y. Miyazaki
Department of Aerospace Engineering, Nihon University, Japan
The authors have been developing a nano-satellite named SPROUT. SPROUT is a 20 x 20 x 22 cm
amateur radio nano-satellite with a mass of 7.1 kg, launched successfully with the Synthetic Aperture
Radar (SAR) satellite ALOS-2 on May 24, 2014. In recent years, space structure, such as a solar sail and a
communication antenna, are becoming larger. But the payload of a rocket has limits on the volume and
the weight. Inflatable membrane structure is very attractive for space structure because it is lightweight
and compact. But it has not been verified sufficiently in space so far. The main mission is to demonstrate
the deployment of a combined membrane structure (1.5m-sided triangular membrane supported with
two inflatable tubes). SPROUT will take the images of the deploying shape of membrane by two
cameras. The image date will be compared with those of the ground experiments and numerical
simulation. Thereby, we estimate reliability of the results of analysis.
In this presentation, the author introduces the initial operation results, operation plan based on the
initial operation results and detailed system of the membrane deployment mechanism.
M onopole antenna (x4)
210
.0[m
m]
Sun sensor (x6)
Camera(x1)
Camera(x2)
1500[mm]
x
y
m]
.8[m
214
Before M embrane Deployment
y
x
M embrane
z
15
00
[m
m
]
220.0[mm]
z
60deg
Solar cell (x30)
I nflatable tube(x2)
After M embrane Deployment
91
Implementation of an On-Board Computer & a Modem into a
Single Subsystem for CubeSat
M.E. Bas1, M.S. Uludag1, I.E. Akyol1 , M.D. Aksulu1,M.Karatas2 and A.R. Aslan2
1
2
ERTEK Space Sys. Co., Istanbul, Turkey
Istanbul Technical University, Istanbul, Turkey
The importance of CubeSats has been increasing dramatically in the last decade due to their short
development time and cost effectiveness. Since there are constraints for the volume and mass of the
CubeSat, it is critical to reduce the space consumed by the mandatory subsystems. The main purpose of
this study is to develop a joint subsystem, which includes an On-Board Computer and a UHF/VHF
Modem which is developed by ERTEK Space Systems Co.; thus, leaving more space for scientific units
and payloads.
BeEagleSat is a 2U CubeSat, which is being developed with the cooperation of various universities and
industrial corporations, within the context of the QB50 Project. The payloads of the BeEagleSat are
MNLP, which is going to be supplied within the context of the QB50 project, and the X-Ray detector,
which is being designed in-house with the cooperation of ITU and Sabanci University.
The OBC part of the joint subsystem, is constituted of an ARM based 32-bit microcontroller, a memory
unit, and failsafe precautions such as; an external watchdog timer and buffers for communication buses.
The OBC will control the other subsystems in accordance with the flight algorithm. The communication
between the OBC and the other subsystems can be established through various communication
protocols (i.e. I2C, SPI, UART).
The Modem part of the joint subsystem is used for data communications and telecommand. SI4463 RF
ICs are responsible for RF side of modem. The received data will be processed and handled by the same
ARM-based microcontroller. The RF parameters (e.g. frequency, modulation type, gain control) are
adjustable through the OBC software via API of SI4463 IC. The used data protocol is compatible with the
AX.25, which is a commonly used by radio amateurs. The modem is also capable of transmitting and
receiving data simultaneously. The data rate will be 9.6kbps for UHF band and 1.2kbps for VHF band.
The joint subsystem will both control the CubeSat subsystems and will be in charge of transmitting and
receiving data from the ground station. By comparison to the currently available, separate Modem and
OBC subsystems, this joint subsystem is more cost & mass effective and space saving.
92
BIRDY: an interplanetary CubeSat to collect radiation data on the way to Mars
with a precursor flight around the Earth in GTO
Boris Segret1, Jordan Vannitsen2, Marco Agnan2, Audrey Porquet3,4,5, Oussema Sleimi2, Jim Lin2, Damien
Boisseau2, Florent Deleflie3,4,5, Jiun-Jih Miau2, Jyh-Ching Juang6, Kaiti Wang7
1
2
Laboratoire d'Etudes Spatiales et d'Instrumentation en Astrophysique, (LESIA), Observatoire de Paris,
Meudon, France
National Cheng Kung University, Department of Aeronautics and Astronautics, Tainan, Taiwan
3
Institut de Mécanique Céleste et de Calcul des Ephémérides (IMCCE) Observatoire de Paris, Paris, France
4
Centre National de la Recherche Scientifique (CNRS), France
5
6
7
Université Pierre et Marie Curie, Paris, France
National Cheng Kung University, Department of Electrical Engineering, Tainan, Taiwan
National Cheng Kung University, Institute of Space and Plasma Sciences, Tainan, Taiwan
BIRDY is a 3-Unit CubeSat that is piggy-backed on a host mission to Mars and jettisoned at the beginning of
the journey. Then it operates in full autonomy: no assistance, no communications but a beacon signal. The
mission profile is a new contribution in Space Weather monitoring and an opportunity to assess the risks in
the manned missions to Mars. It counts energetic particles in the maximum range 1 MeV/nucleon to 1
GeV/nucleon. The ground segment prepares a fine-tuned trajectory to be stored on-board, on the basis of
the planed trajectory of the host mission that provides the main delta-V but not the ideal path. It makes the
CubeSat compatible with almost all missions going to Mars. During the cruise, the CubeSat relies on an
optical planet tracking system to locate itself and on small electrical thrusters to adapt its trajectory and
perform the exact flyby at Mars that permits to come back to the Earth. The science data are collected all
along the journey and only uploaded once in Mars' vicinity to one of the existing Martian orbiters or rovers,
and once at the arrival back to the Earth.
A BIRDY protoflight model is expected to be ready for a precursor flight around the Earth in GTO by 2018 in
order to test the innovative functions of the mission such as the autonomous navigation and
communications subsystems.
More widely than its own scientific mission, BIRDY demonstrates a new way to gather data from distant
locations in the solar system. The project is an educational space mission, essentially leaded and designed
by students from different educational levels in France and in Taiwan.
93
Target Shape Identification for Nanosatellites using Monocular Point Cloud
Techniques
Mark Post, Xiu-Tian Yan
Space Mechatronic Systems Technology (SMeSTech) Laboratory,
Department of Design, Manufacture and Engineering Management
University of Strathclyde, 75 Montrose St, Glasgow, G1 1XJ, UK
Many mission scenarios for nanosatellites and CubeSat hardware have already been created that will
require autonomous target tracking and rendezvous maneuvers in close proximity to other orbiting
objects. While many existing hardware and software designs require the use of rangefinders or
laser-based sensors to identify and track nearby objects, the size and power limitations of a CubeSat
make a simple monocular system greatly preferable, so long as reliable identification can still be
carried out.
This presentation details the development and testing of an embedded algorithm for visually
identifying the shape of a target and tracking its movement over time, which can include rotation
about any axis. A known three-dimensional geometric model is required for use as a reference when
identifying a target. First, feature descriptors implemented in the OpenCV framework are used to
create a sparse point cloud of features from a nearby object. Using structure-from-motion (SfM)
methods, feature points obtained over successive images can be triangulated in three dimensions to
obtain a pose estimate. Statistical shape recognition is then used to identify the object based on
features from available three-dimensional models. While more feature points make the
identification more accurate, more computing power is required, and within the limitations of an
embedded system, the balance of speed and accuracy is evaluated. The algorithm is designed to be
efficient enough for feasible operation using embedded hardware useable on a CubeSat, and can be
used with appropriate hardware for real-time operation. An overview of the algorithm and vision
system design is given, and some initial test results for a simulated orbital rendezvous scenario are
provided for some indication of the performance of these methods. Applications of interest for this
type of algorithm include external monitoring of other spacecraft, robotic capture and docking, and
space debris removal.
94
Posters
95
QBITO, the first CubeSat by Universidad Politécnica de Madrid
I.Barrios1, A. Laverón1 and E. Moreno1
1
Universidad Politécnica de Madrid, Madrid, Spain
QBITO is the first CubeSat developed by Universidad Politécnica de Madrid (UPM). It is a 2U CubeSat and
is one of the satellites that compound the project QB50 led by Von Karman Institute (VKI) in Belgium.
The main task of QBITO will be to operate the Ion Neutral Mass Spectrometer (INMS) that is the primary
payload on-board the CubeSat and that will study the properties of the lower thermosphere. The most
outstanding feature of QBITO is the amount of new in-house developments that are present in the
design. These include the structure subsystem, the Electrical Power Subsystem, the Communications
subsystem and a novel antenna deployment mechanism. These new developments are complemented
with Commercial Off-The-Shelf units in order to reach a robust, yet innovative, architecture.
Apart from the INMS, QBITO will carry three other payloads in order to take advantage of the mission as
much as possible. The first is an experiment that will assess the performance of the n-Octadecane as a
Phase Change Material. It is being developed by the University of Liège. The second experiment, the
Medium Wave Infrared Detector, is developed by the Spanish company New Infrared Technologies and
aims at testing this kind of detectors, whose manufacture process is based on the Vapour Phase
Deposited PbSe technology, in space conditions. Finally, the third additional payload, the Experimental
Software, implements an attitude determination and control software whose algorithms are based on
fuzzy control theory. The purpose is to test the suitability of this kind of control techniques for
spacecraft attitude control applications.
QBITO´s ground segment is composed by four ground stations. Two of them are placed in Madrid at the
Spanish User Support and Operations Center (E-USOC) and Escuela Técnica Superior de Ingenieros de
Telecomunicaciones (ETSIT). The other two are provided by Universidad Nacional de Ingeniería (UNI) in
Lima and Universidad Nacional Autónoma de Méjico (UNAM) that are collaborating with the project.
The CubeSat is currently being manufactured and tested by QBITO team which includes UPM staff,
professors and students and it will be ready for the launch scheduled in January 2016. The team has also
the support of the Spanish company SENER, which provides with technical support along the project and
the access to its environmental test facilities.
96
A survey of CubeSats: Present status and trends
Xiaozhou Zhu, Xin Song, Xiaoqian Chen and Yuzhu Bai1
College of Aerospace Science and Engineering, National University of Defense Technology, P. R. China
Over a decade ago, Professors Bob Twiggs and Jordi Puig-Suari proposed the concept of CubeSats. Since
then, CubeSats have gained comprehensive attentions from academia, industries and space agencies
due to their low construction and launch costs, short research and development cycle, promising
capabilities, and ubiquitous launch providers. As a result, the last decade had witnessed a tremendous
surge in the number of CubeSats launched. And by the time this paper was written, more than two
hundred CubeSats have been sent into space.
In this paper, statistical analysis of CubeSats launched is conducted using data collected from a variety of
sources to throw light upon the present status. Then, examples of a wide range of applications are
provided, including scientific research, technology demonstration, communication, earth observation,
etc. Finally, the potential future issues and development trends are explored.
97
Nano-SSoC: Low cost sun sensor for high accuracy
attitude determination in CubeSats
J. M. Moreno1, J. M. Quero2 and P. Castro1
1
2
Solar MEMS Technologies S.L., Seville (Spain)
Department of Electronic Engineering, University of Seville (Spain)
During the last years applications for CubeSats are rapidly diversifying, with increased use in the future
for Earth observation and remote sensing. These platforms need accurate attitude determination and
control systems (ADCS), therefore those instruments - solar panels, antennas and other hardware - have
to be properly oriented in order to perform their functions.
The Department of Electronic Engineering of the University of Seville and the spin-off company Solar
MEMS have been developing sun sensors in satellite and industrial applications for many years. A new
version specially intended for CubeSats and small platforms, called Nano Sun Sensor on a Chip (NANOSSOC) is a miniaturized two axis sun sensor capable of measuring the incidence angle of a sun ray
accurately in both azimuth and elevation. The sensor consists of four quadrant photodiode fabricated
monolithically in the same crystalline silicon substrate, including a transparent cover glass on the same
silicon die to act as shield to prevent space radiation. The sunlight is guided to the detector through a
window above the sensor, inducing photocurrents on each diode that depend on the angle of incidence.
The novelty of this approach is that SSOC sensing element is based on MEMS technology to achieve high
integrated sensing structures, providing high accuracy, robustness and size and weight reduction.
NANO-SSOC have both analog and digital versions with interfaces fully compatible with most CubeSat
structures and OBCs, with a dimensions around 2,0 x 1,5 x 0,5 cm, a mass close to 5 g and an accuracy
better than 0.5º (3σ) in a 120º FOV. The device integrates electronic circuitry for signal amplification and
conditioning. In addition to this, the digital NANO-SSOC version includes a microcontroller integrated for
selecting, filtering and processing the amplified outputs, directly providing the sunlight incident angles
via UART, SPI or I2C communications interface. A similar version of these sensors have already been
integrated in NANOSAT-1B, launched in 2009, in SEOSAT satellite as a payload and in CEPHEUS CubeSat
mission, which is planned for launch in the first half of 2016.
98
The RIBRAS Software System
W. S. Lisboa, L. R. Hissa, L. R. Amaduro, L. G. L. Moura,
R. A. de Carvalho and C. S. Cordeiro
Instituto Federal Fluminense, Campos dos Goytacazes (Rio de Janeiro), Brasil
The RIBRAS (Brazilian Integrated Satellite Tracking Network) system is a collection of softwares with the
objective of controlling the network of ground stations that will receive the Brazilian satellites data. The
ground stations will work in a synchronized fashion, following a previous generated working plan, to
make sure that the maximum amount of data is collected during satellites’ overflights.
The Software is organized into the master/slave communication model. Each of the Server (master) and
Client (slave) packages has three modules: Scheduler, Distributor and Controller for the server, and Job
Controller, Antenna Controller, and Communication Controller for each client - the ground stations.
On the master side, the Scheduler is responsible for generating an integrated plan for tracking the
satellites of a given constellation during a certain period of time - usually the time between the first and
the last satellites. The Distributor divides the plan into specific jobs that will be dispatched to each
correspondent station. Finally, the Controller will supervise the whole network functioning and is
capable of analysing and re-starting the scheduling process if something goes wrong.
On each station - the slave side - the Job Controller receives the incoming jobs, prepares their execution,
and at the right time executes them. During a job execution, the Antenna Controller and Communication
Controller will control, respectively, the antenna position and radio functioning.
99
VISION: A VIS-NIR atmospheric spectral imager operated from a triple CubeSat
1
D. Fussen, 1D. Pieroux, 1S. Ranvier, 1J. De Keyser and 1P. Cardoen, 1E. Dekemper, 1F. Vanhellemont and
2
H. Saari
1
Belgian Institute for Space Aeronomy, Brussels, Belgium
VTT Microelectronic Systems, Tietotie 3, Espoo, P.O.Box 1000, FI 02044 VTT, Finland
2
The Visible Spectral Imager for Occultation and Nightglow (VISION) is a tuneable spectral imager active
in the visible and near-infrared. It targets primarily the observation of the Earth's atmospheric limb
during orbital Sun occultation. By assessing the radiation absorption in the Chappuis band for different
tangent altitudes, the vertical profile of the ozone is retrieved.
A secondary objective is to measure the deformation of the solar disk so that stratospheric and
mesospheric temperature profiles are retrieved by inversion of the refractive ray-tracing problem.
Finally, occasional full spectral observations of polar auroras are also foreseen.
This miniaturized hyper-spectral imager will be carried as the primary payload of the PICASSO IOD
CubeSat mission and it will be developed by VTT Research Centre in Finland. VISION is innovative in
space applications for remote sensing and it will be based on MEMS-based Fabry-Perot Interferometer
technology. This technology allows the imager to operate over a variable spectral range with a 10 nm
spectral resolution whilst remaining physically compact. Optimisation of the VISION design for the
PICASSO mission is expected to result in stratospheric ozone measurements to 5% accuracy with a
vertical resolution of 2 km after post-processing.
100
FR03 EntrySat
A.Sournac1, J. Chaix1, R.Garcia1, D.Mimoun1
1
ISAE France
Entrysat has for main scientific objective the study of uncontrolled atmospheric re-entry. This project, is
developed by ISAE in collaboration with ONERA and University of Toulouse, is funded by CNES, in the
overall frame of the QB50 project. This nano-satellite is a 3U CubeSat measuring 34*10*10 cm3, similar
to secondary debris produced during the break up of a spacecraft. EntrySat will collect the external and
internal temperatures, pressure, heat flux, attitude variations and drag force of the satellite between
≈150 and 90 km before its destruction in the atmosphere, and transmit them during the re-entry using
the IRIDIUM satellite network. The result will be compared with the computations of MUSIC/FAST, a
new 6-degree of freedom code developed by ONERA to predict the trajectory of space debris.
In order to fulfil the scientific objectives, the satellite will acquire 18 re-entry sensors signals, convert
them and compress them, thanks to an electronic board developed by ISAE students in cooperation with
EREMS. In order to transmit these data every second during the re-entry phase, the satellite will use an
IRIDIUM connection. In order to keep a stable enough attitude during this phase, a simple attitude orbit
and control system using magnetotorquers and an inertial measurement unit (IMU) is developed at ISAE
by students.
A commercial GPS board is also integrated in the satellite into Entry Sat to determine its position and
velocity which are necessary during the re-entry phase. This GPS will also be used to synchronize the onboard clock with the real-time UTC data.
During the orbital phase (≈1 year) EntrySat measurements will be recorded and transmitted through a
more classical “UHF/VHF” connection.
101
Technical aspects of the SNUSAT-1 operation simulator
J. H. Park1, J. Lim1, C. W. Kang1, M. Kim1, Y. D. Kim1, C. G. Park1, H. J. Kim1, S. Kim2 and I.-S. Jeung1
1
Department of Mechanical and Aerospace Engineering, Seoul National University, Korea
2
Department of Biomedical Engineering, Seoul National University, Korea
As one of the development process of SNUSAT-1, a simulator has been put together in order to perform
analysis on operation sequence and satellite state. The fundamentals of the simulator runs on orbital
dynamics including perturbation effects due to geopotential, atmospheric drag, solar radiation pressure,
and third-body, modelled using EGM96 geopotential coefficients, Jacchia77 atmosphere model, and van
Flandern ephemerides model. Fourth order Runge-Kutta method is implemented as the main integrator.
Quasi-inertial reference frame, orbital reference frame and body reference frames are implemented for
attitude simulation, with SNUSAT-1 physical characteristics such as its shape or subsystem elements
represented in vectors within the body frame. Operation analysis includes the operation sequence, such
as solar panel deployment or detumbling and control, communication with ground stations, power
generation and consumption and various states of SNUSAT-1. The analysis is performed by
implementing power model, ADCS models, and time-stamped operation sequence. The work describes
the technical aspects in detail with descriptions of assumptions made for certain models and its
parameters suitable for operation simulation.
102
A Review of Attitude Determination and Control Subsystem of a Nanosatellite
Parth Garg1, Hitesh Agarwal1 and Pinakin M. Padalia1
1
Birla Institute of Technology and Science, Pilani, Rajasthan, India
Advances in highly reliable commercial electronics, miniaturization techniques and materials have
enabled university student teams to design, build and launch nanosatellites. Nowadays nanosatellites
have proved significant as a technology test bed for the Universities. The mission success is critically
dependent on the reliability, consistency and accuracy of spacecraft. The attitude determination and
control system is important to autonomously orient the spacecraft and control the vehicle. The increase
in mission complexity drives the need for a more precise ADCS.
The lessons from previous ADCS designs have proved to be valuable for future mission designs, testing
and integration. The purpose of this paper is to elucidate the process of designing phase of ADCS by
taking help from previous cubesats missions. The paper presents review of ADCS designs of five
successful cubesat missions namely Aausat3, Swisscube, Rax, Delfinext and Cubestar on the basis of
hardware selection of sensors and actuators for pointing as well as stabilization purpose and intricacies
involved in the design electronics. The satellites were chosen on the basis of mission success and testing
of new capabilities in the ADCS design.
The paper helps the beginners to get an insight of the various components of the ADCS and their
interconnections. It helps the experts to review different ADCS designs and come up with the best
combination of sensors, actuators and algorithms for their requirements. It also focuses to help the
reader in evaluating different hardware and software for the ADCS on the basis of power, cost, size,
weight and the harsh space environment. The paper gives a general idea of various stabilization and
pointing methods implemented in previous missions according to the level of control requirements. The
estimation and control algorithms are compared along with their advantages and limitations. With the
detailed system design of the reviewed satellites, the paper also provides their experiences and
modifications made by the designers in their successor satellites. Hence an attempt is being made to
provide the readers with a structured approach in designing ADCS.
103
Simple and robust algorithm for CubeSat attitude estimation
P. Zagórski, A. Tutaj, T. Dziwiński
AGH University of Science and Technology, Kraków, Poland
One of the crucial problems concerning control of artificial Earth's satellites is determination of the
spacecraft spacial orientation. Accurate attitude estimate is required for example for precise
communication antenna pointing. The estimation problem is particularly challenging for very small
satellites (eg. CubeSats), where large and expensive sensors like star trackers are impractical. In the
following paper a computationally inexpensive Quaternion Steepest Descent Attitude Estimator (QSDAE)
algorithm is presented. It is designed to take advantage of minimal number of arguably the simplest
measurements that enable attitude estimation. For this purpose two vector quantities have been
selected: the measurement of the Earth's magnetic field vector and the relative gyroscopic (MEMS)
measurement of satellite angular rate. Both of those measurements can be obtained by cheap,
lightweight and energy-efficient sensors. Additionally estimator does not require knowledge of satellites
inertia tensor, shape or size. The paper presents derivation of the algorithm and realistic computer
model of the environment and measurements. Results of several simulations and computer tests are
included. A procedure of estimator tuning balancing convergence and noise rejection requirements is
proposed.
104
New Star Identification Algorithm Using Ring Projection and Vector Sum
Ki-Duck Kim, Su-Jang Jo, Hyo-Choong Bang
Korea Advanced Institute of Science and Technology, Daejeon 305-701, Republic of Korea
In this paper, a new identification algorithm using ring projection transform (RPT) and vector sum is
proposed for small satellites. The proposed algorithm is one of the pattern recognition methods among
several categories of star identification algorithms. Vector sum, which is created from a ring projection
transform, has rotation invariant characteristics. Also it represents a pattern as very simple imaginary
value. Therefore, star identification can be performed using the captured with reduced amount of
database.
Many researches about ADCS (Attitude Determination and Control System) of small size satellites have
been actively studied in the current satellite area. These small satellites have constraints on using star
trackers because of its size. Limited performance of flight computer causes smaller database capability
and low-speed operation. Miniature star trackers for small satellites require an algorithm, which can
identify star rapidly under restricted database condition. Also, CMOS (Complementary Metal-OxideSemiconductor) active pixel devices are frequently used for miniature star trackers rather than CCD
(Charge-Coupled Device) because of the low power consumption and high update rate. However, the
image Noise of CMOS active pixel devices is relatively larger than CCD images. Consequently, a robust
identification algorithm is important to use CMOS devices for small satellites.
Vector sum using ring projection transform is a well-used method in image pattern matching area.
Calculated vector sum from images becomes a one imaginary value. Only two values, which are its norm
and phase angle, are required in the final comparison step. High identification speed and low amount of
database are expected with this identification algorithm. Identification failure, which can be occurred by
employing the Cartesian coordinates can be avoided with vector sum algorithm because its rotation
invariant characteristics. Using this approach, the robustness for low quality images could be achieved.
105
Technological experiments on-board the URSA MAIOR nanosatellite
L. Arena, F. Piergentili, F. Santoni, B. Betti, F. De Cesare, F. Nasuti, M. Onofri
Sapienza Aerospace Research Center (CRAS), University of Rome “La Sapienza”, Rome (IT)
The Sapienza Aerospace Research Center (CRAS) is involved in the design and manufacturing of the
nanosatellite URSA MAIOR (University of Rome la SApienza Micro Attitude In ORbit testing), scheduled
for launch in January 2016 as one of the Cubesats selected in the framework of the QB50 mission,
leaded by the Von Karman Institute for Fluid Dynamics. In addition to the mNLP QB50 scientific payload,
a drag sail deorbiting system for nanospacecraft and a cold gas microthruster for attitude control will be
tested on board.
The drag sail is made of a special polymeric material. This material works like a memory shape or elastic
material. A sample of this material can be compressed and stored in a small volume. Once the
compression is removed, the material expands again to its original shape. After the deployment, this sail
will use the drag force to deorbit the spacecraft. A first prototype has been realized at Alma Mater
Studiorum – University of Bologna. It passed both vibration and thermo-vacuum tests. The main goal of
the cold gas microthruster experiment (MEMIT – MEMS MIcroThruster) is to test a new integrated
MEMS (Micro Electro Mechanical System) valve-nozzle system. The whole system is designed to fit in
less than a 1/2 U of the CubeSat. The cold gas propellant is nitrogen at ambient temperature. The MEMS
nozzle and valve are manufactured by means of innovative techniques: the present MEMS nozzle has an
axis symmetric geometry and it is controlled by a MEMS valve which works mainly like an
electromagnetic valve. The micropropulsion test consists in providing a constant thrust for a given
amount of time and measuring the angular velocity induced by the thruster on the CubeSat by means of
the gyroscopes.
The electronic control system is designed to survive on-orbit for the entire mission (at least one year). It
coordinates all the telemetry information to be sent to the radio (currents, voltages, temperatures,
earth magnetic field strength, satellite angular velocities and sun vector orientation) and the payloads
operations. It is based on a multi-microcontrollers and FPGAs cold redundancy scheme. This allows both
to enhance the overall satellite reliability and to test on-orbit the endurance of different electronics
technologies (microcontrollers and FPGAs of different brands).
106
The FIRST-S Project: technical challenges
V. Lapeyrère, S. Lacour, L. Gauchet, S. Arroub, R. Gourgues, and G. Perrin
Laboratoire d'Etudes Spatiales et d'Instrumentation en Astrophysique (LESIA),
Observatoire de Paris, France
The FIRST-S project is an astronomical project in the context of exoplanet detection. The goal is to
measure the amount of exozodiacal light scattered by dust around bright nearby stars. To do this we
need high dynamic range (103) at moderate resolution (arcsec).
The proposed instrument is 30 cm baseline stellar interferometer with nulling capabilities based on a
LiNbO3 active optic on a 3U CubeSat. This nulling technique is currently developed in laboratory, and is
suitable for a nanosat application.
The challenging parts of this project are to control the injection of the light in a single-mode fiber with a
accuracy of 1 arcminute and to control the optical path difference between the two arms of the
interferometer at the level of few nanometer, while the CubeSat stability is about 0.1°.
A three stages control solution is proposed. The first stage is the pointing and stability capabilities of the
ADACS system of the CubeSat, reaching 0.1° accuracy. The second stage is to mount the fiber on a three
axis piezo nanopositioner controlled via a position sensing diode. We now reach an accuracy better than
1 arcminute (the star light is properly injected into the fiber) and an OPD accuracy of 1µm. The last stage
is to control the optical phase difference to the nanometer and to scan the null fringe using the active
part of the LiNbO3 component. The interferometer itself is used as an OPD sensor.
With these 3 layers of control we can reach the accuracy in terms of pointing and OPD control.
107
The GOSMOZ Project, an Innovative CubeSat Development Platform
F. Jordan1, J.-M. Jordan1, J. Harris1, B. Chapuis1, J. Iseli1 and J. Selz2
1
2
ELSE SA, Carrouge, Switzerland
EPFL Space Center, Lausanne, Switzerland
Experience has shown that high-quality CubeSats (like SwissCube) are very custom-designed mainly
because of their high system integration. Therefore, they cannot be easily used as platforms for new
payloads. A system-level approach focusing on innovations that change the way of building and
developing CubeSats led to the creation of a new development platform. This project, called GOSMOZ,
not only is a very useful tool for the next generations of CubeSat developers; it also represents the first
family of Swiss-made products available on the CubeSat market.
The feasibility and success of the next generations of CubeSats will increasingly depend on a very good
budget control. Starting CubeSat development from the very beginning is expensive and such cost
quickly becomes a showstopper. With a very efficient development platform at the disposal of
developers, the budget can be better controlled. The first objective of the study was therefore to
determine which innovative technologies or mechanisms would help to drastically reduce the duration
and risks of the development, the integration as well as the testing of a CubeSat. The study shows how it
is achievable by dividing the frame into several slices, inter-connecting the subsystems through the
satellite side-panels with spring-loaded contacts, etc.
The second objective of the study was to figure out how developers can spend more time focusing on
the development of their subsystems instead of trying to solve the recurring problems of subsystems
integration. The answer was found through a smart standardization that provides a high potential for
customization, where it is really needed. The concept can be applied to any CubeSat sizes, up to 12U.
The project’s final objective was to build a website, allowing developers to define their CubeSat online,
to obtain CAD files, mechanical parts, as well as electrical components (connectors, kill-switches and
complete CubeSats subsystems).
108
Development of a sweeping Langmuir probe instrument for monitoring the
upper ionosphere on board a triple CubeSat
S. Ranvier, P. Cardoen, J. De Keyser, D. Pieroux, D. Fussen
1
Belgian Institute for Space Aeronomy, Belgium
A novel Langmuir probe instrument, which will fly on board the Pico-Satellite for Atmospheric and
Space Science Observations (PICASSO), is under development at the Belgian Institute for Space
Aeronomy. PICASSO was initiated to join the QB50 project as scientific in-orbit demonstrator.
The sweeping Langmuir probe (SLP) instrument is designed to measure both plasma density and
electron temperature at an altitude varying from about 400 km up to 700 km from a high inclination
orbit. Therefore, the plasma density is expected to fluctuate over a wide range, from about 1e6/m³
at high latitude and high altitude up to 1e12/m³ at low/mid latitude and low altitude. The electron
temperature is expected to lie between approximately 1000 K and 3000 K.
Given the high inclination of the orbit, the SLP instrument will allow a global monitoring of the
ionosphere with a maximum spatial resolution of the order of 150 m. The main goals are to study 1)
the ionosphere-plasmasphere coupling, 2) the subauroral ionosphere and corresponding
magnetospheric features, 3) auroral structures, 4) polar caps, and 5) ionospheric dynamics via
coordinated observations with EISCAT’s heating radar.
To achieve the scientific objectives described above, the instrument includes four thin cylindrical
probes whose electrical potential is swept in such a way that both plasma density and electron
temperature can be derived.
Along the orbit, the Debye length is expected to vary from a few millimetres up to a few meters. Due
to the tight constraints in terms of mass and volume inherent to pico-satellites, the use of long
booms, which would guarantee that the probes are outside the sheath of the spacecraft (several
Debye lengths away), is not possible. Consequently, the probes might be in the sheath of the
spacecraft in Polar Regions. Extensive modelling and simulations of the sheath effects on the
measured current/voltage characteristics will be performed to ensure an accurate parameter
extraction from the measured data. Another issue implied by the use of a pico-satellite platform for a
Langmuir probe instrument is the limited conducting area of the spacecraft which can lead to
spacecraft charging. In order to avoid this problem, the spacecraft potential is monitored and the
probes are swept in a particular way with limited duty cycle. The resulting measurement data rate is
compatible with the limited telemetry bandwidth available on PICASSO, which will have an S-band
downlink session when it passes over the ground station every few orbits.
109
CubeSat Constellation Design for Zonal Mutual Coverage: Comparison
between Rider Analytical Design and Genetic Algorithm Optimization
I. Meziane-Tani1,2
1
Laboratoire de Télécommunications de Tlemcen, University of Tlemcen, Algeria
Géoazur, Nice Sophia-Antipolis University, CNRS (UMR 7329), Observatoire de la Côte d’Azur, 250
rue Albert Einstein, Sophia Antipolis 06560 Valbonne, France
2
Satellite constellations offer a better spatial diversity and are sometimes the only solution to some
mission requirements. In global coverage case, the design is relatively simple and some models
already exist (Walker and Ballard constellations). However when the coverage becomes more
complicated (regional, zonal), the design is no longer deterministic and several analytical, semianalytical and numerical approaches are proposed in the literature. On the other hand, using
CubeSats instead of convectional large satellites minimizes the development and launch costs in
order to match with regional mission budget.
In this paper, we consider the simple case of a data collection mission where the transmitting station
and the receiving station are both within the same geographical zone. The CubeSats relaying data are
considered without any intersatellite links. First, the Rider analytical model used to find an
approximate constellation to the problem of zonal coverage is evaluated in the case of mutual
coverage. Then, these results are numerically optimised using a genetic algorithm that we
implemented. Finally, through some examples we show that the number of satellites has been
significantly reduced.