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ESTCube-1 In-Orbit Experience and Lessons Learned (Harry Rowe Mimno Award 2016)

Authors:
  • UT Tartu Observatory

Abstract and Figures

ESTCube-1 is a one-unit CubeSat that has been in orbit since May 2013. It was launched to a Sun-synchronous 670 km altitude polar low Earth orbit, and its primary mission objective was to centrifugally deploy a tether as a part of the first in-orbit demonstration of electric solar wind sail (E-sail) technology. The electrical power system, the communication system and the command and data handling system remain fully functional after almost two years in orbit. The camera, developed to image the end-mass of the tether, has taken more than 270 images for camera characterization, for validating the attitude determination system, and for public outreach purposes. The attitude determination accuracy is better than 2°, and the attitude control system is able to spin up the satellite to more than two rotations per second around an axis that suits the E-sail experiment. In this article, we present our in-orbit experience of operating and preparing the satellite for the experiment, as well as lessons learned from development and in-orbit phases.
Content may be subject to copyright.
magazine
Aerospace and Electronic
IEEE
SYSTEMS
August 2015
ISSN 0885-8985
Volume 30 Number 8
12 IEEE A
&
E SYSTEMS MAGAZINE AUGUST 2015
ESTCube-1 In-Orbit Experience and Lessons Learned
Andris Slavinskis
Mihkel Pajusalu
Henri Kuuste
Erik Ilbis
Tõnis Eenmäe
Indrek Sünter
Kaspars Laizans
Hendrik Ehrpais
Paul Liias
Erik Kulu
Jaan Viru
Jaanus Kalde
Urmas Kvell
Johan Kütt
Karlis Zalite
Karoli Kahn
Silver Lätt
Jouni Envall
Petri Toivanen
Jouni Polkko
Pekka Janhunen
Roland Rosta
Taneli Kalvas
Riho Vendt
Viljo Allik
Mart Noorma
Tartu Observatory
Toravere, Estonia
INTRODUCTION
ESTCube-1 is a student satellite project lead by the University
of Tartu, Estonia, and supported by the European Space Agen-
cy (ESA) via Plan for European Cooperating States (PECS).
Development of ESTCube-1 has been a collaborative effort
with many international partners. The satellite is shown on
Figure 1 [1].
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periment [1]–[3]. Implemented according to the one-unit Cube-
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10×10×10 cm and mass of slightly over 1 kg. ESTCube-1 con-
sists of the following subsystems: electrical power system (EPS)
[5]; communication system (COM); command and data handling
system (CDHS) [6]; attitude determination and control system
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payload [11]. All subsystems and payloads were custom built
mostly using commercial off-the-shelf (COTS) components. The
satellite was intended to prepare for and to perform the E-sail
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In-orbit validation.
1. Characterize novel subsystems (EPS, ADCS, and camera).
2. Spin-up the satellite to one rotation per second.
3. Test tether deployment.
4. If deployment successful, charge the tether synchronously
with the satellite spin and measure changes in the spin rate
caused by Coulomb drag interaction between the tether and
the ionospheric plasma.
5. Characterize on-board electron guns.
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cellent platform for educational and in-orbit demonstration (IOD)
projects that are at the same time challenging from the engineer-
ing point of view [12]. The CubeSat standard and the associated
philosophy allow for rapid development [13] and provide the
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eral CubeSat programs have demonstrated how lessons learned
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performed studies of large plasma formations in the ionosphere
Authors’ current address: Tartu Observatory, Space Technol-
ogy, Observatooriumi 1, Tõravere, Tartu county, 00560 Estonia.
E-mail: (andris.slavinskis@estcube.eu). Current addresses for
all authors appear on page 22.
Manuscript received March 5, 2015 and ready for publication
June 15, 2015.
DOI No. 10.1109/MAES.2015.150034.
Review handled by M. Jah.
0885/8985/15/$26.00© 2015 IEEE
Figure 1.
ESTCube-1 satellite before delivering it to the launch provider.
AUGUST 2015 IEEE A
&
E SYSTEMS MAGAZINE 13
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mission to end two months into the mission after it was launched
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mission outcomes and focus on compiling a list of lessons learned
has allowed for the AAUSAT program from Aalborg University
to be successful and continue for more than ten years [19]–[21].
([DPSOHVRIRWKHUVXFFHVVIXO&XEH6DWVHULHVLQFOXGHWKHVDWHOOLWHV
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lin Institute of Technology [24], [25]; the CP CubeSats from Cali-
fornia Polytechnic University [26]; the two DICE satellites from
Utah State University [27]; the Cute series from Tokyo Institute
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Nevertheless, the project has achieved most of its objectives. The
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mission, a need for in-orbit recalibration of attitude determination
sensors, ferromagnetic materials aligning the satellite frame with
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However, from developing all subsystems in-house and operating
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the follow-up missions.
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from the point of view of system engineering, electrical engineer-
ing, mechanical engineering, software engineering, testing and
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IN-ORBIT EXPERIENCE
ESTCube-1 was launched on May 7, 2013 on-board the Vega
rocket by Arianespace. After successful early operations, several
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minimal software functionality to eliminate the risk of activating
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maturely enabling the high voltage supply, unlocking the tether
reel or the tether end-mass.
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functionality: power saving methods, including satellite-wide
timed sleep modes and battery level thresholds for automatically
turning off other subsystems; variety of data logging functions;
a callable timed beacon function for public outreach purposes;
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Similarly to the EPS, the CDHS has been improved by adding
functionality: power saving mode, variety of data logging func-
tions, high time-resolution functions for sensor measurements,
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measurements, as well as attitude determination and control al-
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all calculations take place on the CDHS microcontroller (MCU).
A secondary objective of the ESTCube-1 mission was to take
images of Estonia. Firstly, to validate the camera for this purpose,
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was taken on May 15, 2013. During its lifetime, ESTCube-1 has
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es. These images have been used to characterize the camera and
to validate on-board attitude determination. Due to challenges
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and only at the one-year anniversary was the team able to present
an image of Estonia, Latvia, and a part of Finland (Figure 2). The
most important software updates for the camera were histogram
analysis that allowed automatic detection of the Earth and clouds,
and optimization of power consumption.
Attitude determination sensors were prelaunch calibrated in
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surements. For calibration, statistical methods were used, and at-
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14 IEEE A
&
E SYSTEMS MAGAZINE AUGUST 2015
ESTCube-1 In-Orbit Experience and Lessons Learned
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the system. The accuracy of the system is better than 1.5° [9].
Due to ferromagnetic steel structural components and battery
casings, as well as ferromagnetic nickel anode and cathode of elec-
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with the engineering model and Helmholtz coils in an anechoic
chamber revealed that the residual magnetic moment is larger than
the on-board coils can produce and the direction is roughly diago-
nal from one edge to another. Under stable unactuated conditions
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magnetic moment vector (see Figure 3), which in turn follows the
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tion is not stable and over time the satellite returns to its natural
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was able to reach the spin rate of 360 deg/s.
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by camera nor angular velocity measurements. The most prob-
able reason is that the tether reel is not rotating because either the
rotator is jammed or reel lock deployment has failed (see Section
VIII for more details). To enhance the centrifugal pull force of the
end-mass in an attempt to release the possible mechanical jam,
the spin rate was increased to as high as possible which resulted
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emission-based electron guns, intended to charge up the satellite
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were still tested by powering up the high-voltage source and ap-
plying a potential difference of around 510 V between the electron
gun anode and cathode. Currents going to electron guns measured
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increases the cathode current, indicating that electron guns func-
tion. A voltage of 510 V produced a cathode current of 300 ȝA.
The reliability of the technology still seems to be of concern. One
of the electron guns appears to have disconnected from the power
supply and the functioning one short circuited during tests (after
the successful measurement of the cathode current).
After two years and two weeks of being operational, due to
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ergy-negative mode and consumed the available energy stored in
the batteries to keep operating. Once the batteries were drained,
the satellite did not have enough energy available to be opera-
tional.
SYSTEM ENGINEERING
MODEL PHILOSOPHY
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deliver the satellite on time. On August 2012 the schedule was ac-
celerated by moving the delivery date from May 2013 to January
2013. The decision was a trade-off between engineering risks and
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the risk of components becoming damaged before the launch. In
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reel to turn and break the tether into small pieces, which covered
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solved (solution in Section VIII), in the future, we plan to use a
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(parts of the payload, reaction wheels, thrusters) might not be in-
cluded in all models.
Figure 2.
A composite image showing Estonia, Latvia and a part of Finland taken
on April 23, 2014.
Figure 3.
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The magnetic moment is determined in a laboratory using the engineer-
ing model which did not have electron guns and could be magnetized
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aligned with sides of the satellite.
AUGUST 2015 IEEE A
&
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Slavinskis et al.
STANDARDIZATION
During ESTCube-1 development, each subsystem team was able
to make design decisions independently. Such approach did not
cause any major problems, but we think that all subsystems should
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development tools where applicable, to allow reusability, to save
development time, and to facilitate mobility of team members be-
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XVHG²RQH IRU FRPSXWDWLRQLQWHQVLYH VXEV\VWHPV DQG DQRWKHU IRU
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STANDARDS AND DOCUMENTATION
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Space Standardization (ECSS) can be used as a best practice, sub-
missive following of the ECSS standards introduces too much
overhead for CubeSat projects which usually use agile develop-
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lite. However, the team must use standards and conventions that
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dards (e.g., [33]) have to be followed by CubeSat teams operating
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suggest using web-based documentation tools and/or versioning
and revision control systems. In that case all members can easily
access the newest version (as well as the history of versions) and
maintaining versions is much easier.
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a functionality to log a static set of housekeeping data, but for
in-orbit debugging, dynamic logging of various parameters was
UHTXLUHGDQGLPSOHPHQWHGLQRQHRIWKHVRIWZDUHXSGDWHV
Interface documents must contain detailed descriptions of
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surement units (radians and degrees) between functions imple-
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by updating software. The units must be agreed beforehand but,
as a safety measure for such a risk, a team can introduce correc-
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In addition to the recommendations listed above, we would
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radation, attitude determination and control, as well as payload.
INTEGRATION
The ESTCube-1 team, similar to other CubeSat teams, faced
PDQ\FKDOOHQJHV ZKHQ¿WWLQJ YDULRXVZLUHVDQGFDEOHKDUQHVVHV
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include them in computer-aided design (CAD) mechanical mod-
els. Integration of subsystems and components should be prac-
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gest maintaining as fully functional as possible a prototype of the
satellite that contains the latest subsystems to test prototypes of
new subsystems. In this case, many problems could be detected
right when the new revision of the component is inserted into the
satellite assembly. Another option is assembling as complete a
model of the satellite as possible on a periodical basis and per-
forming conformity tests.
In which order the side panels attach to the satellite frame
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side panel before all connections under that side panel have been
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panels as independent of each other as possible to reduce the ef-
fect of these problems.
To remember to integrate all components, they should be laid
out on a table. A simple but effective way to ensure a success-
ful integration is to make a checklist of all components and pro-
cesses. Development of the checklist should start early and all
subsystems should be involved.
Prior to the integration in the cleanroom all the components
KDYHWREHFOHDQHGWKRURXJKO\WR PHHWWKHVWDQGDUGVUHTXLUHGE\
the launch provider and the middlemen, and to ensure that the
components like solar panels and lenses will not become con-
taminated. Contamination can accumulate on lenses, causing arti-
facts on images, and on solar panels, reducing the amount of solar
photons that can reach solar cells (therefore, effectively reducing
WKHHI¿FLHQF\8VLQJ SURWHFWLYH ¿OPV ZKLOH LQWHJUDWLQJDQGUH-
moving them before the launch can help to avoid contamination.
In the case of ESTCube-1, the satellite was successfully in-
tegrated and some of the suggestions listed above were followed
but by fully following them the integration process can be opti-
PL]HGIXUWKHUDQGPDGHPRUHWLPHHI¿FLHQW
ELECTRICAL ENGINEERING
COMMERCIAL OFF-THE-SHELF COMPONENTS
The electronics on-board ESTCube-1 were assembled solely from
COTS components, a market which is developing rapidly, and
WKHUHIRUHKLJKSHUIRUPDQFHFRPSRQHQWVFDQEHREWDLQHGTXLFNO\
and at low cost. To ensure reliability, automotive or industrial-
grade components were used, where possible, and several redun-
dancy measures were applied to assure that a component failure
would not jeopardize the mission and also several tests were per-
formed (see Section VII).
7KHLQÀLJKWH[SHULHQFHVKRZVWKDWWKLVDSSURDFKZDV DVXF-
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sor and a failed memory) did not cause any larger problems due
to redundant counterparts of components. Applying redundant
16 IEEE A
&
E SYSTEMS MAGAZINE AUGUST 2015
ESTCube-1 In-Orbit Experience and Lessons Learned
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ciency due to power electronics components working in parallel
and sharing the load.
DATA CONNECTIONS WITHIN THE SATELLITE
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bus standards, both between the components of a single subsys-
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ferent subsystems, we used the universal asynchronous receiver/
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according to our in-house developed internal communication
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(e.g., analog-to-digital converters (ADCs), magnetometers, input/
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grated circuit (I2C) and serial peripheral interface (SPI) buses [1].
The main challenges arose from cases when the same com-
munications bus was shared between several components, espe-
cially when the systems connected could be powered on and off
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operational due to the current supplied through a communications
bus, even if the component itself is not powered through its power
pins. Also, a single unpowered device on a bus can drain enough
current to make the whole bus inoperable when communicating
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of a switch for disconnecting unpowered devices from buses or
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latter might not always achieve the results needed.
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two electrical connections, one for transferring data and the other
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the state machine behind I2C communications can malfunction,
leading to the loss of communication capability with the compo-
nent. Therefore, it should be possible to separately power off I2C
devices to reset their internal state. This is not a problem with the
SPI bus. It also has happened that communicating with a single
device using the I2C bus causes other devices on the shared bus
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oscillator chip for beacon systematically malfunctioned when an
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arise from the fact that on an I2C bus, a single data line is oper-
ated both by the bus master and the bus slave, making level con-
version complicated.
All in all, we would suggest refraining from using I2C in sat-
ellites, especially for critical communications. If an I2C bus is
shared between several components, it is advisable to implement
some form of a chip select functionality and have an option to
separately power off or reset components. As another note, SPI
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MEMORY
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for nonvolatile storage of system-critical data because the un-
derlying technology is highly radiation tolerant. However, one
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ample, secure digital (SD) cards. This allows using third-party
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than developing them in-house. However, in the case of memory
devices with integrated controllers, abrupt power loss becomes
an issue.
Parallel memory devices should be used where applicable.
Although the current consumption of parallel memory is higher
when compared with serial memory devices, parallel interface
provides greater performance and makes them easier to address.
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mission success.
ELECTRICAL POWER
Producing and distributing electrical power proved to be a chal-
lenging task both while designing the system and during opera-
tions in orbit. For more details about the design, see [5].
In the design phase, one of the largest challenges was imple-
menting redundancy measures, especially due to the large number
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regulators were duplicated within the EPS in a hot redundant con-
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to monitor redundant systems. Fortunately, no power component
failures were detected during the operations period. The power
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critically analyzing the need for redundancy, during shorter mis-
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lite and before the launch; this time period can be easily over-
looked in the design process. During this time the satellite has to
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tery energy consumption when the satellite is inside the satellite
deployer, since the satellite might remain in that state for months
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the satellite in orbit (overdraining the batteries during this period
might also cause irreversible damage).
As mentioned in Section II, one serious problem we encoun-
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panel degradation will take place during every satellite mission
and this often determines the mission lifetime. Therefore, we sug-
gest a highly granular power distribution system in which com-
ponents and subsystems can be powered off independently, con-
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automatic battery voltage thresholds, which caused automatic
subsystem turn-off when the power level became critical. This
AUGUST 2015 IEEE A
&
E SYSTEMS MAGAZINE 17
Slavinskis et al.
system can be developed to automatically achieve power posi-
tivity, even in case of communication problems. To further save
power, we also used timer-based sleep modes, in which only the
EPS was powered. Still, great care must be taken so that the sys-
WHPH[LWVWKHVHPRGHVUHOLDEO\HYHQLQWKHFDVHRIPHPRU\RYHU-
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used control areas before and after critical memory sections, in
addition to checksums of these sections. It is also a good idea to
implement an automatic system to hard-reset the whole satellite
if the satellite has not been successfully communicated with for
VRPHWLPH:HXVHGDKRXUWLPHUIRUWKLV
An important conclusion from automatic power saving fea-
tures is that all critical data should be kept in nonvolatile mem-
ories. In the case of ESTCube-1, we lost some camera images,
IRUH[DPSOHGXHWRWKHLUQRQYRODWLOHVWRUDJHV\VWHP6KRUWWLPH
power failures might also happen for other reasons, including ra-
diation effects and software errors.
All in all, the power system implementation managed to pro-
vide enough power for the satellite to reduce the problem of solar
panel degradation from a mission stopper to a minor inconvenience.
OTHER
The ESTCube-1 CDHS has two cold redundant MCUs that are
VHOHFWHG E\ WKH (36 7R UHGXFH LQWHUVXEV\VWHP FRPSOH[LW\WKH
on-board computer can have its own low-power radiation-tolerant
processor for critical administrative tasks as well as for switching
the main MCUs.
,QWUDVXEV\VWHPEXVHVVKRXOGQRWEHH[SRVHGWRRWKHUVXEV\V-
tems to avoid possible compatibility issues that would affect the
performance of components within a subsystem.
MECHANICAL ENGINEERING
MAIN STRUCTURE
A mono-block aluminum structure was used on ESTCube-1 be-
cause it is lightweight and it makes it easier to achieve the re-
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tegration, we will not use a mono-block structure in the future.
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for the main structure compared with the one suggested by the
CubeSat standard (aluminum alloy 6061 or 7075) [4] because
it was easier to order in Europe. Changes in the main structure
material did not cause any problems, but the last minute change
from titanium to steel bolts introduced ferromagnetic material on
board. Suppliers and products should be secured early to avoid
late changes.
In a perfect case, the launcher should be known during the de-
YHORSPHQWSKDVHRIWKHVWUXFWXUHEHFDXVHWKHUHTXLUHGWROHUDQFHV
change from launcher to launcher.
8QLTXHPDWHULDOVVKRXOGEHDYRLGHGWRKDYHDFKDQFHWRUHSUR-
duce mechanical structures after the launch.
Apart from the ferromagnetic bolts, all ESTCube-1 issues re-
garding the main structure are minor.
SOLAR PANELS
Solar panel cover glass should be used to avoid rapid degradation
of solar cells. In the case of ESTCube-1, we did not use cover
JODVVVLQFHLWVLQKRXVHDSSOLFDWLRQLVFRPSOH[DQGLWUHGXFHVWKH
EHJLQQLQJRIOLIH HI¿FLHQF\ RI VRODU SDQHOV :H DOVR XQGHUHVWL-
PDWHGWKHH[WHQWRIGHJUDGDWLRQGXULQJWKHWLPHUHTXLUHGWRFRP-
plete the mission. Lack of solar panel cover glass was likely the
main cause of the rapid solar panel degradation on ESTCube-1,
and in hindsight we strongly suggest using cover glass, even for
shorter missions, and especially on polar orbits (higher amount of
trapped particles encountered).
SUN SENSORS
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WLYLW\RIWKHVHQVRUPDVN²LQWHUQDOVXUIDFHVVKRXOGEHDEVRUELQJ
black to avoid stray light on position sensitive devices. In a per-
fect case, the mechanical design of the sensor mask would not
allow the incident light to illuminate internal surfaces. In the case
of ESTCube-1, the aluminum mask was anodized black and the
GHVLJQFDQEHLPSURYHGWRDYRLGXQZDQWHGUHÀHFWLRQVLQVLGHWKH
mask.
CAMERA
The basic aluminum structure of the ESTCube-1 camera lens
HQFORVXUHSURYLGHVD VXI¿FLHQWDPRXQWRIUDGLDWLRQSURWHFWLRQLQ
DORZ(DUWK RUELW 5DGLDWLRQ DIIHFWV WKHFDPHUD5$0ZKLFK LV
located right behind a 1 mm thick side panel [10]. The effect can
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sensor can produce. For ESTCube-1, these effects are not critical
but, if they would be, memory devices could be protected with
shielding.
The imaging sensor is also prone to radiation effects. Perma-
QHQWO\GDPDJHGKRWSL[HOVFDQEHDYRLGHGZLWKWKHKHOSRIDVKXW-
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VSHFWUDOUDGLDWLRQ¿OPVDQGUREXVW¿OWHUVFDQEHXVHG
MOMENT OF INERTIA
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FULWLFDOWRGHWHUPLQHWKHDWWLWXGHSUHFLVHO\ZKHQSHUIRUPLQJ¿QH
attitude maneuvers and especially when high spin rate maneuvers
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control and an attitude estimator with a prediction step is used
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tant because in the prediction step the attitude is propagated using
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QDOYDOXHVRIWKHLQHUWLDPDWUL[ZHUHHVWLPDWHGDQDO\]LQJLQRUELW
PHDVXUHPHQWVRIWKHVSLQSODQH6XFKDSSURDFKSURYLGHGUHTXLUHG
results for attitude determination for low spin rates. However, an
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to grow when the angular velocity increased.
18 IEEE A
&
E SYSTEMS MAGAZINE AUGUST 2015
ESTCube-1 In-Orbit Experience and Lessons Learned
CONNECTORS
On ESTCube-1, the system bus is based on the PC/104+ standard
connector that has 4×SLQVDQGLWVVWLIIQHVVPDNHVLWGLI¿FXOWWR
assemble or disassemble the satellite. The placement and stiffness
of the connectors must be planned thoroughly and coordinated
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to minimize mechanical tensions during integration or disintegra-
tion. As the standard connector heights were not properly taken
into account in the structure design, the pins of some connectors
had to be trimmed. However, challenges with connectors did not
cause any major problems.
SOFTWARE ENGINEERING
OPERATING SYSTEM
In order to minimize the computational overhead and memory foot-
print of the on-board software, a lightweight real-time operating sys-
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important reason to use an operating system was the need for task
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vices that cannot be accessed directly and due to a limited amount of
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the parts that have changed), which would have been useful. If pos-
sible, we suggest using an operating system that provides most of the
QHHGHGIXQFWLRQDOLW\IRUH[DPSOHDIRUPRIHPEHGGHG/LQX[
SOFTWARE UPDATES
A large proportion of ESTCube-1 software was written after the
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orbit software updates and the mission is delayed, increasing a risk
RIVDWHOOLWHIDLOXUHEHIRUHSHUIRUPLQJDOOWKHSODQQHGH[SHULPHQWV
However, we think that functionality of in-orbit software updates
of all active subsystems is critical for a CubeSat mission, especial-
O\IRUWHDPVZLWKRXWSULRUH[SHULHQFH7KDWIXQFWLRQDOLW\FDQPRVW
importantly, save the mission and it also allows using the satellite
IRURWKHUSXUSRVHVWKDQLQLWLDOO\SODQQHG:KHQLPSOHPHQWLQJVXS-
port for software updates, the bootloader must be designed to keep
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IHULQWKHFDVHRI(67&XEHSDJHE\SDJHXSORDGVDQGYHUL¿FD-
tion by pagemaps and checksum have served well.
OTHER
The camera was designed following a principle of using as few
components as possible, which has worked well to provide a small,
simple, modular, and independent camera. However, it should be
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The CDHS is able to log a single command response to a
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UDPHWHUV WR ¿OHV VLPXOWDQHRXVO\ FDQ HDVH WKH SUHSDUDWLRQ DQG
compression of the telemetry.
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reset the MCU) can be used to optimize power consumption of
a system.
A central communication bus is preferred so that subsystems
would be able to communicate with each other without forward-
ing packets through each other.
Developing on-board algorithms in C can save time spent on
SRUWLQJ)RUH[DPSOH$'&6IXQFWLRQVZULWWHQLQ&FDQEHWHVWHG
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directly used in on-board software.
Downlink data rate could be improved further if the COM were
able to buffer several packets in its memory and transmit them in a
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ing a forward error correction coding on the downlink channel.
An obvious but important lesson learned is to document the
code and keep user manuals up to date.
Lessons learned presented in this subsection did not cause any
major problems but can make development and operations more
HI¿FLHQW
TESTING AND MEASUREMENTS
CALIBRATION AND CHARACTERIZATION
All on-board sensors have to be calibrated and characterized to
gain measurement reliability. It should include as many test cases
as possible. Planning of tests has to start early in the project be-
FDXVHVRSKLVWLFDWHGWHVWEHQFKHVPLJKWEHUHTXLUHG)RUH[DPSOH
DWWLWXGHVHQVRUVVKRXOGEHURWDWHGDURXQGDOOD[HVVLPXOWDQHRXVO\
to develop reliable calibration curves.
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LQJWHPSHUDWXUH VHQVRUV LQFORVHSUR[LPLW\WR RWKHU VHQVRUVDQG
performing temperature-calibration for all on-board sensors can
improve the accuracy of other sensor measurements remarkably.
Combining laboratory calibration with in-orbit calibration
might give the best result because not all cases can be tested in a
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)RUH[DPSOHHQGWRHQG$'&6WHVWLQJPLJKWDOZD\VKDYHVRPH
limitations.
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mal vacuum and vibration tests on a subsystem level before the
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perfect case, sensors must be calibrated under conditions that are
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Sun sensor could measure an incidence angle of light while being
placed in a thermal vacuum chamber.
To decrease analog sensor uncertainty, a temperature-com-
pensated reference voltage should be measured on board.
In the case of ESTCube-1, some sensors were well calibrated
before the launch but on multiple occasions in-orbit measure-
ments had to be used to recalibrate them. For attitude determina-
tion sensors, we were lacking test benches that would provide
the needed variety of tests. More temperature and voltage sensors
will be used on board upcoming satellites.
Another important aspect is the timing of measurements. In
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AUGUST 2015 IEEE A
&
E SYSTEMS MAGAZINE 19
Slavinskis et al.
HDVLO\KDSSHQ WKDWFXUUHQWDQG YROWDJHYDOXHVWDNHQ LQVHTXHQFH
actually correspond to different power states, making calculations
based on multiple sensor readings problematic. A solution would
be an independent telemetry system with synchronized input buf-
fers to be certain that all measurements correspond to the same
moment of time. Filtering can also be used to reduce this problem.
INFANT MORTALITY
Infant mortality is an early component failure caused by not test-
LQJVXI¿FLHQWO\ZHDULQJDVHQVRUEHIRUHWKHODXQFK:HKDYHH[-
perienced failure of one of four hot redundant gyroscopic sensors
soon after the launch and one of the two cold redundant MCUs of
WKH&'+6VXIIHUHGGDPDJHWRWKHLQWHUQDOÀDVKMXVWWKUHHPRQWKV
DIWHUWKHODXQFK7ZRRXWRIWKUHH63,EXVÀDVKPHPRU\GHYLFHV
on the ESTCube-1 engineering model stopped working a few
weeks after the integration. Flight and spare components should
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PLWLJDWHWKHULVNDVZHOO:HFRQVLGHUWKLVLVVXHHVSHFLDOO\LPSRU-
tant with COTS components.
MAGNETISM
Having ferromagnetic materials on-board the satellite has caused
WKHELJJHVWFKDOOHQJHLQSUHSDULQJWKHVDWHOOLWHIRUWKHH[SHULPHQW
It took more than half a year to partly characterize the magnetic
properties of the satellite using the engineering model, to fully
characterize the motion of the satellite in orbit, and to iteratively
improve and test attitude controllers. Nevertheless, the spin-up
PDQHXYHUFRXOGQRW EH SHUIRUPHG DV SODQQHGIRUWKH(VDLOH[-
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GHSOR\ZLWKRXWVLJQL¿FDQWGHÀHFWLRQDJDLQVWWKHVDWHOOLWHVLGHVHH
)LJXUH:HVWURQJO\VXJJHVWFKDUDFWHUL]LQJPDJQHWLFSURSHU-
WLHV RI ÀLJKW FRPSRQHQWV DQG WKH PRGHO SULRU WR WKH ODXQFK LQ
WKHFDVH DQDWWLWXGH FRQWURODFWLYHRUSDVVLYHLV UHTXLUHGLQ WKH
magnetosphere of the Earth. Note that this issue affects not only
VDWHOOLWHVWKDWXVHPDJQHWRUTXHUV
OTHER
A practice of early prototyping should be combined with regu-
lar subsystem-level functional tests followed by early integration
tests (starting with electrical/software and later adding mechani-
cal tests) to develop a well-functioning and reliable system.
'HGLFDWHG ERDUGV IRU HDUO\ WHVWV FDQ EH FRQVLGHUHG IRU H[-
ample, to test and perform preliminary characterization of a va-
riety of sensors from which the best ones can be chosen for the
mission.
To make debugging and diagnostics easier, test-ports can be
OHIWRQDÀLJKWPRGHODQGDXQLYHUVDOVHULDOEXV86%FRQQHFWRU
can be used at least until an engineering model is prepared.
The power budget must account for the degradation of solar
cells and batteries.
In the case of ESTCube-1, we incrementally learned lessons,
listed in this subsection, and applied them to our activities on the
go.
ELECTRIC SOLAR WIND SAIL (E-SAIL) PAYLOAD
The ESTCube-1 tether payload consists of a piezoelectric motor
driven reel, 25 ȝm and 50 ȝm wires forming a 15 m long tether,
an end-mass of the tether, a high voltage source to charge the
tether, and a slip ring to connect the high voltage supply to the
WHWKHU5HHOLQJRIWKHWHWKHULVPRQLWRUHGE\WDNLQJLPDJHVRIWKH
HQGPDVV%RWKWKHHQGPDVVDQGWKHUHHODUH¿[HGZLWKGHGLFDWHG
locks that use burn wires [11].
:HFDUULHGRXWWKH WHWKHUGHSOR\PHQW WHVWLQ RUELWDQGLWZDV
not successful. Since the payload design suffers from a lack of
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failed is not known. Some of the future design improvements for
the E-sail payload are sensors to detect whether locks have de-
SOR\HGLIWKHUHHO LVWXUQLQJ DQGLI WKHHQGPDVV LVPRYLQJ%\
having the camera inside the tether enclosure, it will be possible
to monitor the end-mass even before deployment. In the case of
(67&XEHWKHHQGPDVVZRXOGDSSHDULQWKH¿HOGRIYLHZRQO\
after tether deployment of a few centimeters. To improve end-
mass monitoring even further, a light-emitting diode (LED) is
VXJJHVWHGWREHDGGHGQHDUWKHHQGPDVVHQFORVXUHIRUSUR[LPLW\
imaging. Such LED would allow imaging of the most critical pe-
riod of tether deployment without depending on sunlight and/or
DWWLWXGH7KH(67&XEHFDPHUDLVOLPLWHGWRVWRULQJDPD[LPXP
of four images. More memory would allow monitoring deploy-
ment in detail. Nonvolatile memory should be used to avoid los-
LQJH[SHULPHQWGDWDLQWKHFDVHRIDUHVHW
As described in Section II, the reel started to turn during the
TXDOL¿FDWLRQ YLEUDWLRQ WHVW DQG WKH WHWKHU JRW EURNHQ$V D ODWH
design change, a reel lock was introduced. To avoid late design
FKDQJHVVXEV\VWHPVKDYHWREHTXDOL¿HGVHSDUDWHO\EHIRUHLQWH-
gration. In the case of the payload, vibration tests were envisaged
EXWGXHWRDODFNRIUHTXLUHGWHVWVSHFL¿FDWLRQVWKH\FRXOGQRWEH
accomplished.
To couple the tether rotation to the spacecraft spin, the tether
mechanical attachment point should reside as far from the space-
craft center of mass as possible. The tether then resembles a rotat-
ing pendulum (rod attached to a spinning plate) maintaining its
nominal orientation with respect to the spacecraft body. However,
given the dimensions of the tether reel and one-unit CubeSat, this
LVKDUGWRDFFRPSOLVK7KXVWKH(67&XEHWHWKHU ZDVH[SHFWHG
WRRVFLOODWHLQDFRQHRIDERXWGH¿QHGURXJKO\E\WKHGLPHQ-
VLRQV RI WKH HQGPDVV RSHQLQJ +RZHYHU LI WKH WHWKHU GHÀHFWV
more than about 20°, it would touch the conductive side panel.
This would lead to wearing of the tether and even a short circuit.
To avoid these risks, an additional grommet should be placed to
the side panel opening. The grommet must be of antistatic mate-
rial to avoid the triple junction with the plasma, high voltage, and
nonconducting material. For ESTCube-1, tether movement also
decreases the chance to image the end-mass as the end-mass is in
WKH¿HOGRIYLHZRQO\ZKHQWKHWHWKHULVQHDULWVQRPLQDORULHQWD-
tion, normal of the satellite side panel.
Another part of the E-sail payload is the high voltage (HV)
supply system and the electron guns. On the HV supply side, the
PRVWFRQFHSWXDOO\GLI¿FXOWSUREOHPZDVPDQDJLQJZKLFKSDUWVRI
the payload and the satellite are referenced to the HV source and
20 IEEE A
&
E SYSTEMS MAGAZINE AUGUST 2015
ESTCube-1 In-Orbit Experience and Lessons Learned
WRWKHUHTXLUHGHOHFWURQLFV7HVWVSHUIRUPHGLQRUELWVKRZWKDWWKH
HV supply board is operational. Developing the telemetry col-
lection system of the HV board was a challenge due to the fact
WKDWWKHHOHFWULFDOJURXQGOHYHORIWKHV\VWHPÀRDWHGZLWKUHVSHFW
to the satellite ground, since ADCs were referenced to the satel-
lite ground. In the future, we suggest putting telemetry collection
HOHFWURQLFVFRPSOHWHO\LQWKHÀRDWLQJJURXQGVLGHDQGRQO\XVLQJ
digital communication lines to interface them. This should also
make the calibration of the system easier.
In the case of the electron guns, reliability remains the major
concern. In our tests, one gun did not work at all (open circuit)
and the other short circuited during tests (although it was con-
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the reliability would have been to test the whole HV and electron
gun system together in the laboratory before the launch. In the
time frame of ESTCube-1, this was not possible; also it would
KDYHUHTXLUHGDUDWKHUFRPSOH[YDFXXPFKDPEHUVHWXS$QRWKHU
possible cause for problems was the fact that the electron guns
were not covered during deployment and their surface might have
gotten contaminated before their use (even a small particle could
short circuit the system). This could be mitigated by using protec-
tive covers that would be removed in orbit or by increasing the
distance between the cathode and the anode.
MANAGEMENT
TEAM LEADING
Having a visionary leading the team is key to successfully car-
rying out a technically challenging and long project, especially
a project where team members must be regularly motivated by
RWKHUWKDQ¿QDQFLDOPHDQV7HDPPHPEHUVPXVWNQRZDQGDFFHSW
that an ultimate measure of success might come after years of
development, after launching, and after successfully operating a
satellite. However, it is up to leaders and the management team to
GH¿QHIUHTXHQWPLOHVWRQHV
ADVISORY
Another key to success is having professional advisers to su-
pervise and to review student work. In the case of ESTCube-1,
SURIHVVLRQDODGYLFH ZDVUHFHLYHG LQWKH¿HOGV RIHOHFWULFDO HQJL-
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and measurement sciences. Advisers from the amateur radio com-
munity helped to avoid many problems with practical commu-
nication and ground station set-up. Similarly, a system engineer
should be a professional with a wide knowledge of involved en-
JLQHHULQJ¿HOGV
MANAGEMENT
Management of the ESTCube-1 project has been successful in
FDUU\LQJRXWVRPHRILWVIXQFWLRQV)RUH[DPSOHWKHWHDPKDVDF-
cess to various tools and services like the team collaboration soft-
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versioning and revision control system SVN, Google services
(Mail, Documents, Hangouts, Calendar), the chat and video con-
ference system HipChat, a remote desktop computer with access
WR VSHFLDOL]HG VRIWZDUH 0$7/$% 6LPXOLQN 6ROLG:RUNV WKH
3&%VRIWZDUH($*/(HWF+RZHYHUZHKDYHLGHQWL¿HGWKDWWKH
management team cannot consist of people whose main duties are
other than management of the ESTCube-1 project. At least one
person should respond to issues on a daily basis, especially during
critical moments such as integration, testing, prelaunch servicing,
and launch. Special attention should be paid to procurement han-
dling. For management to be responsive, it should be supported
¿QDQFLDOO\
The team should acknowledge the possibility of failures and
EHSUHSDUHGIRUWKHP0DQDJHPHQWFDQSOD\DVLJQL¿FDQWUROHLQ
leading the team to this understanding. Management must take
into account that students can contribute only a limited amount of
time to the project and that they can leave at any time.
In case the project schedule is accelerated, it should be agreed
XSRQZLWKH[WHUQDOSDUWQHUVDQGVXEFRQWUDFWRUV7KHPDQDJHPHQW
team has to follow the development progress on a weekly basis
in order to successfully schedule milestones and ultimately the
launch.
:HWKLQNWKDWFKRRVLQJDFKDOOHQJLQJVFLHQWL¿FPLVVLRQLVEHW-
ter than a simple one (e.g., optical camera being the main pay-
load) in an educational project. In that case, the project outcomes
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motivation. In the ESTCube team, some members have decided
to continue their professional careers with the E-sail.
TEAM
A student team should be motivated and open. Students and the
team will gain the most if members are able to work on various
subsystems and different types of tasks, including leading, man-
agement, article writing, and outreach. Such approach will also
help to avoid alienation between subsystem teams and teams in
different geographical locations.
From the early stages, subsystem teams should discuss re-
TXLUHPHQWVDVZHOODVGHVLJQFKRLFHVDQGPRVWLPSRUWDQWO\LQWHU-
IDFHVIRUH[DPSOHWRQRWFDXVHGLVFUHSDQF\EHWZHHQWUDQVPLWWLQJ
and receiving interfaces.
The team should have a clear understanding of the importance
of the work and the priorities of its subtasks. Everybody in the
WHDPVKRXOGXQGHUVWDQGDQGVKRXOGEHDEOHWRH[SODLQWKHPLVVLRQ
DQGLWVUHTXLUHPHQWV)LJXUH
2SHQLQJDOOTXHVWLRQVIRUDGLVFXVVLRQFDQOHDGWREHWWHUGHFL-
sions, can contribute to team building and can serve as an infor-
mative media for updating on current progress and future plans.
)RU H[DPSOH DOO PHPEHUV VKRXOG EH LQYROYHG LQ FKRRVLQJ WKH
launch provider and in satellite operations.
3HUVRQDO FRQÀLFWV VKRXOG EH VROYHG ZLWKRXW KHVLWDWLRQ DQG
with the help of peers and leaders.
CONCLUSIONS
In this article, we presented an overview of the ESTCube-1 in-
RUELWH[SHULHQFH DQGGLVFXVVWKH OHVVRQVOHDUQHG$IWHUXSGDWLQJ
AUGUST 2015 IEEE A
&
E SYSTEMS MAGAZINE 21
Slavinskis et al.
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for four problems.
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However, the amount of the produced power was enough to pro-
ceed with the mission. In the future, the problem can be solved by
using solar panel cover glass.
Second, a need to recalibrate attitude determination sensors
in orbit. After recalibrating the sensors, debugging software, and
¿QHWXQLQJ WKH .DOPDQ ¿OWHUWKH DWWLWXGH GHWHUPLQDWLRQ V\VWHP
ZDV SUHSDUHG IRU DWWLWXGH FRQWURO PDQHXYHUV DQG WKH (VDLO H[-
periment. In the future, all sensors have to be calibrated before
the launch better than was done on ESTCube-1, as well as full
integration of the system has to be tested on the ground. However,
in-orbit calibration methods can serve as a backup or can be used
WR¿QHWXQHFRUUHFWLRQSDUDPHWHUV
Third, ferromagnetic materials used on-board aligning the
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possible to prepare the satellite for tether deployment. In the fu-
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ization of on-board materials.
Fourth, the inability to deploy the tether made it impossible
WRPHDVXUHWKH(VDLOIRUFH:KLOHIRU(67&XEHLWZDVQRWSRV-
VLEOHWRH[DFWO\GHWHUPLQHZKDWFDXVHGWKHSUREOHPZHZHUHDEOH
to identify design improvements, some of which have already
been implemented on the Aalto-1 satellite [34]. The tether de-
ployment system has to be thoroughly tested and it has to have
means to detect which part is not working (e.g., locks, the reel, or
the tether is broken).
In addition to learning from the four major problems, we have
GLVFXVVHGRWKHUOHVVRQVOHDUQHGLQWKH¿HOGVRIV\VWHPHQJLQHHU-
ing, electrical engineering, mechanical engineering, software
engineering, testing and measurements, as well as management.
Since the satellite delivery schedule was accelerated, the mis-
sion encountered delays. After launching the satellite, only pre-
liminary validation was feasible. New software updates allowed
to fully validate the on-board systems, provide full functionality,
and optimize power consumption.
Lessons learned, discussed in this article, have already been,
DQGFRQWLQXHWR EH DSSOLHG WR VXEVHTXHQW PLVVLRQVLQWKH(67-
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lite size, the target audience for this article is the nanosatellite
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consider standards like the ECSS highly useful. However, they
are not fully compatible with agile development methods that
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:HWKLQNWKDWIRU,2'PLVVLRQVWKDWQDQRVDWHOOLWHVDUHRIWHQXVHG
IRUWKHVWDQGDUGVFDQ EHPDGHORRVHUWRNHHSWKHFRVWHI¿FLHQF\
and short development time.
For teams that are developing satellite series for IODs and
IRUHGXFDWLRQDOSURSRVHVZHHQFRXUDJHXVLQJWKHSKLORVRSK\³À\
HDUO\À\RIWHQ´6XFKSKLORVRSK\HQDEOHVUDSLGWHFKQRORJ\GH-
YHORSPHQWIROORZHGE\LQRUELWWHVWV:KLOHLW LQFUHDVHVWKHULVN
the team can learn from mistakes and unsuccessful missions to
TXLFNO\GHYHORSVHTXHQWLDOVDWHOOLWHV6RPHRIWKHOHVVRQVFDQEH
learned only by launching and operating satellites. Fly early &
À\RIWHQHPSOR\VWKHFRVWHI¿FLHQF\RIQDQRVDWHOOLWHVDQG&276
components. Moreover, launching a satellite soon after freezing
the design allows utilization of the latest developments in the
COTS market.
ACKNOWLEDGMENTS
The authors would like to thank everybody who has contributed
to the development of ESTCube-1. The European Space Agency
and the Ministry of Economic Affairs and Communication of
Estonia have supported ESTCube-1 via the ESA PECS project
“Technology demonstration for space debris mitigation and elec-
WULFSURSXOVLRQ RQ (67&XEH VWXGHQWVDWHOOLWH´:HZRXOG OLNH
to thank all institutions that contributed to ESTCube-1 develop-
ment. The research by Andris Slavinskis was supported by the
European Social Fund’s Doctoral Studies and the International-
L]DWLRQ3URJUDP'R5D7KHUHVHDUFKRQVRIWZDUHHQJLQHHULQJZDV
VXSSRUWHGE\WKH(XURSHDQ5HJLRQDO'HYHORSPHQW)XQGDQGWKH
Investment and Development Agency of Latvia via the Latvian
(OHFWURQLFDQG2SWLFDO(TXLSPHQW&RPSHWHQFH&HQWUHLQ3URGXF-
tion Sector (agreement no. L-KC-11-0006) project number 2.9
³5HFHLYLQJYDOLGDWLRQDQGERRWORDGLQJV\VWHPRIVPDOOVDWHOOLWH
software as an enabler for safe and reliable satellite development
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[21] 1LHOVHQ-'DQG/DUVHQ-$([SHULHQFHVDQGOHVVRQVOHDUQHGGXU-
ing the launch and early orbit phase of AAUSAT-3. In WK(XURSHDQ
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$,$$868&RQIHUHQFHRQ6PDOO6DWHOOLWHV, 2011.
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[31] Kleshch, V. I., Smolnikova, E. A., Orekhov, A. S., Kalvas, T., Tarvain-
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[34] Khurshid, O., Tikka, T., Praks, J., and Hallikainen, M. Accommodat-
LQJWKH SODVPDEUDNH H[SHULPHQWRQERDUG WKH$DOWRVDWHOOLWHPro-
ceedings of the Estonian Academy of Sciences, Vol. 63(2S) (2014),
±
A. Slavinskis, H. Kuuste, T. Eenmäe, I. Sünter, K. Laizans, E.
Kulu, U. Kvell, J. Kütt, K. Zalite, S. Lätt, R. Vendt, V. Allik and
M. Noorma are with Tartu Observatory, Department of Space
Technology, Observatooriumi 1, 61602, Tõravere, Estonia, e-
mail: andris.slavinskis@estcube.eu). A. Slavinskis, M. Pajusalu,
H. Kuuste, E. Ilbis, T. Eenmäe, K. Zalite, H. Ehrpais, J. Viru, J.
Kalde, U. Kvell, K. Kahn, S. Lätt, P. Janhunen and M. Noorma
are with the University of Tartu, Institute of Physics, Ravila
14C, 50411, Tartu, Estonia. A. Slavinskis, J. Envall, P. Toivanen,
J. Polkko and P. Janhunen are with the Finnish Meteorologi-
cal Institute, Erik Palménin aukio 1, P.O. Box 503, FI-00101,
Helsinki, Finland. A. Slavinskis, I. Sünter, U. Kvell, J. Kütt, and
R. Vendt are also with SIA Robotiem, Paula Lejin¸ a iela 5-70,
LV-1029, Rı¯ga, Latvia. P. Liias is with Tallinn University of Tech-
nology, Ehitajate tee 5, 19086, Tallinn, Estonia. K. Zalite is with
Engineering Research Institute Ventspils International Radio
Astronomy Centre of Ventspils University College, Inženieru
101a, LV-3601, Ventspils, Latvia. R. Rosta is with the German
Aerospace Center (DLR), Robert Hooke 7, 28359, Bremen, Ger-
many. T. Kalvas is with the University of Jyväskylä, Department
of Physics, P.O. Box 35 (YFL), FI-40014, Jyväskylä, Finland.
... Another payload candidate, a radiation monitoring device, later called RADMON, was proposed by a team from University of Turku and University of Helsinki. The third payload candidate, selected by the study, was e-sail experiment device, EPB, which was already in development at Finnish Meteorological Institute (FMI) for ESTCube-1 CubeSat mission [28,29]. In the original feasibility study, a vibration monitoring system was also proposed. ...
... The CubeSat platform was selected because it provided affordable access to space and also available commercial subsystems for inexperienced team. The payloads were designed concurrently with the satellite platform, AaSI and RADMON were entirely new designs whereas EPB development was already started for ESTCube-1 satellite [28]. ...
Preprint
Full-text available
The design, integration, testing, and launch of the first Finnish satellite Aalto-1 is briefly presented in this paper. Aalto-1, a three-unit CubeSat, launched into Sun-synchronous polar orbit at an altitude of approximately 500 km, is operational since June 2017. It carries three experimental payloads: Aalto Spectral Imager (AaSI), Radiation Monitor (RADMON), and Electrostatic Plasma Brake (EPB). AaSI is a hyperspectral imager in visible and near-infrared (NIR) wavelength bands, RADMON is an energetic particle detector and EPB is a de-orbiting technology demonstration payload. The platform was designed to accommodate multiple payloads while ensuring sufficient data, power, radio, mechanical and electrical interfaces. The design strategy of platform and payload subsystems consists of in-house development and commercial subsystems. The CubeSat Assembly, Integration & Test (AIT) followed Flatsat -- Engineering-Qualification Model (EQM) -- Flight Model (FM) model philosophy for qualification and acceptance. The paper briefly describes the design approach of platform and payload subsystems, their integration and test campaigns, and spacecraft launch. The paper also describes the ground segment & services that were developed by the Aalto-1 team.
... The spacecraft was launched on May 7, 2013 onboard a Vega rocket, but unfortunately a failure of the tether unreel mechanism occurred during the launch phase, and the tether experiment did not take place. Later, ground tests showed that the piezoelectric motor (see Section 2.2) was damaged by the launch phase vibrations [159], so that the failure was not due to an intrinsic flaw of the plasma brake technology, but may be solved with some spacecraft design improvements [160]. ...
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The Electric Solar Wind Sail (E-sail) is an innovative propellantless propulsion system conceived by Pekka Janhunen in 2004 for use in interplanetary space. An E-sail consists of a network of electrically charged tethers maintained at a high voltage level by an electron emitter. The electrostatic field surrounding the E-sail extracts momentum from the incoming solar wind ions, thus giving rise to the generation of a continuous thrust. In a geocentric context, the same physical principle is also exploited by the plasma brake, a promising option for reducing the decay time of satellites in low Earth orbits after the end of their operational life. This paper discusses the scientific advances of both E-sail and plasma brake concepts from their first design to the current state of the art. A general description of the E-sail architecture is first presented with particular emphasis on the proposed tether deployment mechanisms and thermo-structural analyses that have been carried out over the recent years. The working principle of an E-sail is then illustrated and the evolution of the thrust and torque vector models is retraced to emphasize the subsequent refinements that these models have encountered. The dynamic behavior of an E-sail is also analyzed by illustrating the mathematical tools that have been proposed and developed for both orbital dynamics and attitude control. A particular effort is devoted to reviewing the numerous mission scenarios that have been studied to date. In fact, the extensive literature about E-sail-based mission scenarios demonstrates the versatility of such an innovative propulsion system in an interplanetary framework. Credit is given to the very recent studies on environmental uncertainties, which highlight the importance of using suitable control strategies for the compensation of solar wind fluctuations. Finally, the applications of the plasma brake are thoroughly reviewed.
... 2) Calibration of ESTCube-1 In-orbit Data: ESTCube-1 attitude is largely dominated by the torque from residual magnetic moment of the spacecraft, resulting in its magnetometer reading mostly pointing in one side of the spacecraft. The measurement was taken over several measurement campaigns in May 2014, where several attitude control maneuvers were conducted, resulting in better data set for this calibration method [17], [24], [27]. The sensors sampling period varies from 400 to 700 ms, with a total of 2744 measurements. ...
... The charging of the tether can be accomplished by a separate electron gun, or by utilizing the satellites conducting body as the current, thereby balancing the electron gathering surface. In this paper, a microtether is defined as a tether that has average mass of less than 200 milligrams (mg) per length of one meter (200 mg/m) Plasma brake deployment has been attempted twice in orbit, by the Estonian ESTcube-1 in 2013 and the Finnish Aalto-1 launched in 2017 [5] [6]. Both attemps to deploy the microtether failed. ...
Conference Paper
Full-text available
We present a satellite deorbiting simulator with focus on the novel deorbiting technology of the plasma brake, which utilizes the Coulomb interaction between a charged tether and plasma flow to generate thrust. The plasma brake force is tied to atmospheric variables by plasma density while the aero drag force is related to atmospheric density. These parameters vary significantly: from high and low solar activity due to the 11 year solar cycle, to seasonal changes (Russell-McPherron effect) as well as between day and night due to solar illumination. Space environment variables are obtained from NRLMSISE and IRI2016 atmosphere and ionosphere simulators respectively. The atmosphere is separated into 10 km altitude segments. Results show the simulator to give quite realistic estimates on Coulomb drag deorbit. The plasma brake is a very efficient deorbiting tool at altitudes of over 500 km, while at lower orbits atmospheric drag dominates.
... In the Ncube 1U CubeSat, EPS-6 is implemented with only one MPPT converter by connecting all the PV panels in parallel and on the load-side POL converters are used to feed the different loads [70]. The ESTCube-1 CubeSat which has performed first in-orbit electric solar wind sail experiment also utilized EPS-6 with redundant components to ensure that component failures would not jeopardize the mission [59]. The in-flight results showed that component failures did not cause any major problems and system efficiency is improved due to parallel operating converters. ...
Article
CubeSats have been popular for space research due to lower cost, faster development, and easier deployment. The electrical power system (EPS) is one of the significant subsystem for the CubeSat since it handles power generation, energy storage, and power distribution to all other subsystems. Therefore, the design of EPS becomes crucial for successful CubeSat mission, wherein the first step is the selection of EPS architecture. The main objective of this paper is to present an extensive review of all the conventional and emerging EPS architectures of CubeSats. A total of seventeen categories of CubeSat EPS architectures have been identified, classified, and the operational aspects of these architectures are presented in addition to a qualitative comparison. This study is expected to provide a useful reference guide for all the researchers and developers working in the area of CubeSats EPS. Also, some of the potential research topics are provided to further exploration and innovation for the CubeSat EPS.
... This approach is aimed at simplifying the FTC structure and enhancing the attitude control system efficiency for space missions. Based on the experience of developing and operating ESTCube-1 [31], [32], ESTCube-2 continues in-orbit demonstration of the plasma brake and the electric solar wind sail (E-sail), commonly known as the Coulomb Drag Propulsion (CDP) system [33]. CDP experiments require centrifugal tether deployment [34]. ...
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Chapter
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