Aalto-1, multi-payload CubeSat: design, integration and launch
J. Praks, M. Rizwan Mughal, R. Vainio, P. Janhunen, J. Envall, P. Oleynik, A. Näsilä, H. Leppinen, P. Niemelä, A. Slavinskis, J. Gieseler, P. Toivanen, T. Tikka, T. Peltola, A. Bosser, G. Schwarzkopf, N. Jovanovic, B. Riwanto, A. Kestilä, A. Punkkinen, R. Punkkinen, H.-P. Hedman, T. Säntti, J.-O. Lill, J.M.K. Slotte, H. Kettunen, A. Virtanen
AAalto-1, multi-payload CubeSat: design, integration andlaunch
J. Praks a , M. Rizwan Mughal , R. Vainio b , P. Janhunen c , J. Envall c ,P. Oleynik b , A. N¨asil¨a d , H. Leppinen a , P. Niemel¨a a , A. Slavinskis a,e ,J. Gieseler b , P. Toivanen c , T. Tikka a , T. Peltola a , A. Bosser a ,G. Schwarzkopf a , N. Jovanovic a , B. Riwanto a , A. Kestil¨a c , A. Punkkinen b ,R. Punkkinen f , H.-P. Hedman f , T. S¨antti f , J.-O. Lill g , J.M.K. Slotte h ,H. Kettunen i , A. Virtanen i a Department of Electronics and Nanoengineering, Aalto University School of ElectricalEngineering, 02150 Espoo, Finland b Department of Physics and Astronomy, 20014 University of Turku, Finland c Finnish Meteorological Institute, Space and Earth Observation Centre, Helsinki, Finland d VTT Technical Research Centre of Finland Ltd, Espoo, Finland e Tartu Observatory, University of Tartu, Observatooriumi 1, 61602 T˜oravere, Estonia f Department of Future Technologies, 20014 University of Turku, Finland g Accelerator Laboratory, Turku PET Centre, ˚Abo Akademi University, 20500 Turku,Finland h Physics, Faculty of Science and Technology, ˚Abo Akademi University, 20500 Turku,Finland i Department of Physics, P.O.Box 35, 40014 University of Jyvaskyla, Finland
Abstract
The design, integration, testing and launch of the first Finnish satel-lite Aalto-1 is briefly presented in this paper. Aalto-1, a three-unit Cube-Sat, launched into Sun-synchronous polar orbit at an altitude of approxi-mately 500 km, is operational since June 2017. It carries three experimentalpayloads: Aalto Spectral Imager (AaSI), Radiation Monitor (RADMON)and Electrostatic Plasma Brake (EPB). AaSI is a hyperspectral imager invisible and near-infrared (NIR) wavelength bands, RADMON is an ener-getic particle detector and EPB is a de-orbiting technology demonstrationpayload. The platform was designed to accommodate multiple payloads M. Rizwan Mughal is also associated with Electrical Engineering Department, Insti-tute of Space Technology, Islamabad, Pakistan, Correspondence: [email protected]
Preprint submitted to Acta Astronautica January 27, 2021 a r X i v : . [ phy s i c s . s p ace - ph ] J a n hile ensuring sufficient data, power, radio, mechanical and electrical in-terfaces. The design strategy of platform and payload subsystems consists ofin-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 accep-tance.The paper briefly describes the design approach of platform and payloadsubsystems, their integration and test campaigns and spacecraft launch. Thepaper also describes the ground segment & services that were developed byAalto-1 team.
Keywords:
Aalto-1, CubeSat, hyperspectral, radiation, Aalto SpectralImager, Radiation Monitor, Electrostatic Plasma Brake
1. Introduction
Nowadays, there is an increased interest towards small satellite missionsdue to advances in Commercial Off-The-Shelf (COTS) technology miniatur-ization. Traditionally, the classification of small satellites is only based ontheir mass but the CubeSat standard also takes into consideration the vol-ume [1]. Over the past decade, the applications of small satellites in generaland CubeSats in particular have increased manifold due to the availabilityof low-cost design, testing and launch possibilities [2, 3, 4, 5]. Initially per-ceived for training and educational activities, the applications of CubeSatshave expanded in vast application areas in the past few years [6]. Exampleapplication areas include remote sensing, Earth observation, disaster man-agement, science, astronomy, space weather and technology demonstrationetc. [7, 8, 9, 10, 11, 12, 13].The abundant availability of COTS components with faster developmentcycles has led to the NewSpace movement [14]. This approach has led thetransformation of CubeSat missions from educational and technology demon-stration to real missions with potentially risky but higher commercial andscience return [15, 16, 17]. A large number of commercial applications usingCubeSats have evolved in the past few years with a promising future scopeof commercial applications [2, 3, 4, 5]. Until now, more than one thousandCubeSats have been launched into space [18]. However, the forecast suggestsan exponential increase in nanosatellite launches every year [19].There has also been great advancement in the technology development2or nano and microsatellites [20]. A number of innovative platforms havebeen designed and demonstrated in space [21, 22, 23, 24]. Due to tech-nology miniaturization, the capability of CubeSat platforms has been everincreasing [25, 26]. The current CubeSat missions are able to demonstrateinnovative platforms with high power generation, precise attitude pointingand higher data downlink capabilities with potential to compete with theirbigger satellite counterparts.During the past decade, the worldwide trend of the first satellite by eachuniversity or Small & Medium Enterprises (SME) has been designing andlaunching relatively less complex single-unit (1U) CubeSat for capabilitydemonstration. In contrast, we at Aalto university followed a more challeng-ing approach, i.e., designing a multi-payload CubeSat with student teams.The mission objective was to build and launch a spacecraft with focus on sci-ence, imaging and de-orbiting technology demonstration while also providinghands-on educational training. This paper presents detailed design aspectsof the Aalto-1 CubeSat with a capability description of payloads and theplatform to accomplish the mission objectives. The in-orbit demonstrationand lessons learned are presented in an accompanying paper [27].This paper is organized as follows: Section 2 briefly introduces missionobjectives and requirements, Section 3 presents the mission design, projectimplementation and educational outcomes, section 4 presents space segmentdesign and implementation, section 5 presents all payloads: their specifi-cations and designs, Section 6 introduces design approach of the platformsubsystems, Section 7 presents the integration & testing, section 8 focuseson ground segment and Section 9 concludes the paper.
2. Mission objectives
The Aalto-1 satellite project was initiated from Aalto University stu-dent’s aspiration to make the first satellite mission in Finland. The idea wassupported by teachers and developed during a special assignment in SpaceTechnology course in 2010 spring semester in the form of feasibility study ofthe satellite. The goal of the course was set to develop a realistic satelliteconcept which should be possible to implement (at least partly) by students.It was required that the main payload should be developed in Finland andit should be connected to Aalto University curriculum. During the feasibil-ity study this was translated to the goal to build first Finnish satellite withEarth Observation payload. 3or the university, the main driver for the Aalto-1 project was to providehands-on education in space engineering, science and entrepreneurship, whiletaking advantage of the NewSpace movement [14, 2, 3, 4, 5] and harness theenthusiasm of building the first national satellite. It was envisioned, that inaddition to satellite development, students will also learn to work with expe-rienced space scientists and develop connections to industrial partners.Themission was largely financed and led by Aalto University and integrated toAalto space technology curriculum.Despite a main goal of building, launching and operating first nationalsatellite, the proposed payload selection introduced complex technology demon-stration and science goals. The first feasibility study built the satellite con-cept around four payload candidates and established a consortium for build-ing a 3U Cubesat. The study also derived the main mission requirements.As an outcome of the feasibility study, the satellite mission and platformwas to be developed by Aalto University students and payloads were to becontributed by partner organisations.The main payload candidate was a spectral Earth Observation imager,AaSI, based on technology developed by VTT Technical Research Centre ofFinland (VTT). This further led to wider spectral device offering for spaceapplications by VTT. Another payload candidate, a radiation monitoringdevice, later called RADMON, was proposed by a team from University ofTurku and University of Helsinki. The third payload candidate, selected bythe study, was e-sail experiment device, EPB, which was already in devel-opment at Finnish Meteorological Institute (FMI) for ESTCube-1 CubeSatmission [28, 29]. In the original feasibility study, a vibration monitoring sys-tem was also proposed. However, the idea was later abandoned as impracticalfor a rather monolithic nanosatellite.Neither of the selected payloads had flight heritage. Moreover, AaSI andEPB main technology was never demonstrated in space before for proposedpurpose. Earth Observation with tunable Fabry–P´erot Interferometer (FPI)was a novel concept and also deorbiting a satellite with electrostatic force byusing a tether was never attempted before. This provided a technical chal-lenge and scientific novelty for the project. The project consortium, whichconsisted of Aalto University, University of Turku, University of Helsinki,VTT and Finnish Meteorological Institute, decided to build a multi-payloadmission. In a retrospective, one can say that this decision elongated theproject significantly and enforced also several compromises in the design dueto contradicting requirements. It took slightly over five years from the first4dea to FM completion. The overarching Aalto-1 mission objective was tobuild a satellite to carry out in-orbit demonstration of AaSI, RADMON andEPB experiments, each of them with specific mission objectives.AaSI’s main objective was to demonstrate the operation of a tunable FPI-based spectral imager for Earth Observation in the space environment. TheFPI technology developed at VTT allowed to build for the first time freelytunable spectral EO camera to nanosatellite form factor. As a minimum,the instrument was required to take wavelength calibration measurementsand record a spectrum of at least six wavelengths of a cloud-free land. Fora full demonstration, the instrument was required to take measurements toinvestigate wavelength stability, thermal effects and long-term degradationof filters, optics, the sensor and other components along with demonstratingvarious operation modes.RADMON’s main objective was to operate and calibrate a CubeSat-compatible radiation detector which registers protons in nine energy channelswith threshold energies of 10 – 40 MeV and electrons in five energy channelswith threshold energies of 1.5 – 12 MeV.EPB’s main objective was to deploy a tether and then charge it to esti-mate the force exerted by the Coulomb drag between the tether’s electric fieldand the Earth’s ionosphere, as well as to demonstrate de-orbiting by keepingthe tether charged for an extended period of time. This novel propulsionconcept was (and currently still is) never demonstrated in space. By nowtwo launched nanosatellites, ESTCube-1 and Aalto-1 have made attemptsto deploy this system. However, in the near future AuroraSat-1, Foresail-1and ESTCube-2 are heading towards similar goals using the same technology[30].Detailed mission requirements kept developing along the project and werenot documented in detail. Therefore, it can be said, that the mission wastechnology driven as it often happens in CubeSat missions. However, theproper feasibility study in the very beginning and well established consortiumhelped to keep the focus on results.The finally launched satellite followed closely the original plan of payloadsand functionality, but seriously underestimated requirements due to manyconstraints including time and resources.5 . Mission design and implementation
In order to satisfy the payload in-orbit demonstration requirements, Aalto-1 was required to be launched to a polar orbit with an altitude of at least500 km. A polar orbit provides sufficient conditions to estimate the Coulombdrag force [31] and allows for RADMON to measure at various latitudes, in-cluding the South Atlantic Anomaly. Polar orbit also allows coverage inFinland and provides good opportunities for Earth Observation.The attitude requirements were set by all payloads, but dominated byEPB requirements.The lower limit of altitude was required by EPB. In loweraltitudes, the atmospheric drag might dominate the de-orbiting impact whichmakes electrostatic drag estimation difficult. The highest altitude limit wasset by 25-year orbital decay requirement for space debris mitigation. TheEPB experiment requires spinning satellite of hundreds of degrees per sec-ond in order to provide centrifugal force for tether deployment [32]. Theangular momentum was to be provided in steps: spin up the satellite, de-ploy the tether, spin up again, etc. AaSI requires nadir pointing duringimage acquisition and RADMON requires attitude knowledge, but the re-quirement was not critical. Another notable requirement for the mission wassurface conductivity requirement by EPB to keep spacecraft potential duringColoumb drag experiment. The satellite was designed for two years in theorbit, which was estimated as sufficient time to carry out all experiments.The mission design in terms of energy and thermal budget was flexible, as itwas decided that payload duty cycles can be adjusted in-orbit according tothe need.The satellite operation from Aalto University was one of the key mis-sion requirements. For this purpose, the ground segment was developed.The ground station includes UHV, VHF and S-band steerable antennas andassociated transceivers. The mission operation software was designed andimplemented by Aalto students.The product tree of Aalto-1 mission with ground segment, space segmentand launch segment description is shown in Fig 1.
After successful feasibility study in spring 2010, the satellite project wasquickly funded and supported by Aalto MIDE (Multidisciplinary Instituteof Digitalisation and Energy)[33]. The project was also formally organizedby establishing posts for project responsible professor, project coordinator,6 latformPayload
Aalto-1 Ground segmentSpacecraft
EPS
RADMON
Aalto University GS
EPB
OBC TT&C
Solar Panels
Power management
Power distribution UHF transceiver
ADCS
S band transmitter ADSRadio amateur GSs
STRUCT
ACS
Star tracker / Sun sensors
Gyroscope / Magnetometer
GPS
Magnetorquer
AaSI EPB Electronguns Batteries Stack connector
Launcher
CubeSat deployer
OBCSW
Reaction wheels
GPSantenna
Stack interface
UHF antennaS band antenna
Thermal
Figure 1: Aalto-1 product tree
Steering Group, and Scientific Advisory Board. Under the official projectumbrella, student groups developed their own organization for building thesatellite and ground segment. Thematic student teams were in most casesoriented on single subsystem development or a single topic. The qualityassurance was maintained as a separate independent branch as it is practicedin bigger satellite projects. During the semester, student teams had weeklymeetings and decisions were made by team-leader meetings.Thanks to the available funding, it was possible to hire few doctoralstudents, provide summer trainee positions and occasional master thesis po-sitions with salary (usual practice in Finnish Universities). The doctoral stu-dents formed a backbone of the project which helped accumulate the salientknowledge. Many subsystems were developed as a master thesis project.Payload teams formed separate project structure in their home organi-zations and their team leaders were part of the Scientific Advisory Board.7atellite bus and payload developments were financially independent and ap-plied for funds independently. Several satellite project students made theirmaster thesis with payload team.The project schedule was built to mimic larger space projects, where themain project phases were separated by milestone reviews. The PreliminaryDesign Review was arranged in November 2011, Critical Design Review inMay 2013, Test Readiness Review in May 2015 and Flight Readiness Reviewin January 2016. Several smaller reviews were arranged along the project.The flight model of the satellite was delivered to Netherlands in May 2016.Review panels were assembled from space technology professionals andCubeSat team members from other universities. Both, documentation basedreview format (in the beginning of the project) and presentation based reviewformat (towards the end of the project) were used.Flatsat, EQM and FM model policy with fast iterative development modelwas implemented in the project. A single Flight Qualification Model (FQM)approach was considered in the beginning of the project, but it proved to beimpractical. The students were inexperienced and learned most efficientlyby making prototypes and hardware versions, therefore rapid iterations andfrequent hardware models proved to be more efficient than waterfall design.The main challenge for the project was to find and keep the knowledgein the team during multi year project. Student teams were volatile and doc-umentation often incomplete, despite a requirement for documentation toretrieve study credits. The fact that key persons were hired and commit-ted to the project, helped the project to continue. Constant support by theuniversity and project organization was also highly important. The projectwas also well aligned with university goals: it was able to produce degrees,research papers and positive publicity. Aalto University procured and fi-nanced, along with few sponsoring partners, also the satellite launch, thefirst in Finland.
The student work was incorporated to student’s individual studies mainlyvia special assignments, bachelor and master thesis projects and also as apart of doctoral studies. The main challenge was to align project needs andproject documentation with teaching and outcome assessment in situationwhere most of the work was done in groups.During the project evolved a documentation and reporting approachwhere project documentation was used for grading and individual contri-8utions were assessed by self evaluation and peer reviews. Assessment wasdone on the basis of provided snapshot of the evolving documentation andit was required that the documentation was available for entire project. Thefinal grade was assigned by supervising professor [34]. Far more that 100students contributed in the design and development of the satellite. How-ever, the contribution varied from single semester participation in meetingsto several years of design and implementation. Around 12 Master level and28 Bachelor level thesis were conducted in the satellite design and develop-ment activity during the coarse of the project. By now, also three doctoraldissertations are defended based mainly on Aalto-1 satellite related topics[35, 36, 37] and several are still on the way. Additionally more than 10 Mas-ter level thesis were conducted at partner institutes related to the design anddevelopment of payloads. The outcomes and results were also published inmany scientific conferences and journals in the field. The project gathereda lot of media attention which led to the awareness of space technology andsmall satellites in vast areas [38].Many of the Aalto-1 students became space engineers and scientists atpartner institutions. A subgroup of Aalto-1 students established the ICEYEcompany, which builds operates a fleet of Synthetic-Aperture Radar (SAR)satellites. Another group formed Reaktor Space Lab company, which spe-cializes on nanosatellite missions.
4. Space Segment design and implementation
The feasibility study and preliminary design analysis proposed 3U Cube-Sat platform to carry out the mission. The CubeSat platform was selectedbecause it provided affordable access to space and also available commercialsubsystems for inexperienced team. The payloads were designed concurrentlywith the satellite platform, AaSI and RADMON were entirely new designswhereas EPB development was already started for ESTCube-1 satellite [28].The 3U satellite platform was designed and manufactured mainly by stu-dents of Aalto University. However, in early stage of the project it wasdecided that electrical power system and attitude system should be procuredfrom commercial provider. The main reason for that was the reliability con-cern of fresh designs.The satellite design, as shown in Fig. 2, features 3U CubeSat body, 3-axisstabilization, body mounted solar panels, deployable UHF antennas, several9ameras and openings for payloads. Electronics of the satellite is accom-modated in two electronic stacks, connected by cabling. The Long Stackfeatures all the avionics and AaSI payload. The Short Stack accommodatesRADMON and EPB. The main reason for this separation was the designdecision to align the EPB reel motor rotation axis with satellite rotation axisin spinning mode.
Figure 2: The structure and subsystems of Aalto-1 satellite. The highlighted subsystemsare: 1) Radiation Monitor (RADMON), 2) Electrostatic Plasma Brake (EPB) 3) GlobalPositioning System’s (GPS’s) antenna and stack interface board, 4) Attitude Determina-tion and Control System (ADCS), 5) GPS and S-band radio, 6) Aalto Spectral Imager(AaSI), 7) Electrical Power System (EPS), 8) On-Board Computer (OBC), 9) Ultra HighFrequency (UHF) radios, 10) solar panels, 11) electron guns for EPB, 12) S-band antenna,13) debug connector, and 14) UHF antennas.
Figure 3: A block diagram of digital, RF and power interfaces
An overview of the Aalto-1 power, data and Radio Frequency (RF) inter-faces is presented in Fig. 3. The power interface provides regulated voltagelevels (3.3 V, 5 V and 12 V) to the satellite avionics and payloads. Severaldigital interfaces including Inter Integrated Communication (I C), Serial Pe-ripheral Interface (SPI) and Universal Asynchronous Receiver Transmitter(UART) have been implemented which are controlled by the OBC.11 . Payloads
The instrument design technique, mass, volume, electrical and mechanicalinterfaces and key design challenges of the Aalto-1 payloads is described indetail in this and subsequent sections.
The RADMON instrument [39] is a compact low-power radiation monitor.It has envelope dimensions of about 4 × ×
10 cm , a mass of 360 g and a powerconsumption of 920 mW. The spacecraft supplies both +5 V and +12 V tothe instrument. The instrument consists of a detector assembly inside abrass casing, a signal processing board, a digital board, and an electricalpower board. Three boards are connected by a 52-pin internal bus runningthrough all of the boards (see Fig. 4(a)). The instrument is integrated inthe short stack of the satellite with another bus connector as well as withfour spacers placed in the corners of the PCB stack. The bus connector alsoprovides the electrical interface to the satellite.The detector unit consists of a rectangular 2.1 × × silicon de-tector and a 10 × ×
10 mm CsI(Tl) scintillation detector that are enclosedby the brass casing determining the acceptance aperture (see Fig. 4(b)). Thecasing has an aluminum entrance window that protects the detector stackfrom low-energy charged particles and photons. The scintillator has a thinpolytetrafluoroethylene (PTFE) wrapping on five sides and has a readoutphotodiode on the sixth side. We have used a Hamamatsu S3590-08 PINsilicon photodiode with dimensions of 10 ×
10 mm and a depletion thicknessof about 0.3 mm. The silicon detector has a biased guard ring and a floatingone. The passive area of the silicon detector extends to about 0.7 mm aroundthe active spot. Two detectors produce electrical signals for a standard Δ E– E analysis aimed at the determination of particle species and the energydeposited in the detector. A coincidence logic prevents the registration ofparticles coming from outside the aperture and bremsstrahlung X-rays gen-erated in the brass container.The aluminum window sets thresholds for electron detection at about1 MeV and for proton detection at about 10 MeV. The brass case becomestransparent for protons at about 55 MeV, approximately at the same energyas protons incident through the aperture start to penetrate the scintillator.RADMON registers protons in nine energy channels with threshold energiesof 10 – 40 MeV and electrons in five energy channels with threshold energies12 a) The RADMON radiation monitor assembly.
Silicon detectorCsI(Tl) detectorPhotodiodereadout (b) The detector unit.Figure 4: The assembly compounds three printed circuit boards and a brass containerwith a detector unit inside. An aluminum entrance window in front of the brass containercovers the detector unit. A scintillator with a readout photodiode and a silicon detectorare housed within the brass container. of 1.5 – 12 MeV. The detailed analysis of the instrument response to electronsand protons is described in [40]. The data rate can be adjusted by changingthe polling frequency of the instrument. Nominally, science data is collectedevery 15 seconds and housekeeping data every 60 seconds. This gives a datarate of about 25.4 kBytes per hour, including the packet overhead.Testing and ground calibrations of RADMON were performed using ra-dioactive sources and a proton beam from the MGC-20 cyclotron at the ˚AboAkademi University, Turku, Finland. The maximum beam energy availablein the cyclotron was about 17 MeV. The beam was scattered at about 60degrees from a thin tantalum foil to lower the beam intensity and achieve alow-enough flux for the calibrations. The proton beam energy was step-wisedecreased by adding absorbers between the foil and the detector. This setupallowed to successfully calibrate the instrument for the low-energy protonresponse, and Geant4 [41, 42] simulations were used to extend the proton13esponse over the full energy range. The electron response was monitoredutilizing beta particles from different radioactive decay sources.Radiation tolerance of RADMON electronics has been tested in the RA-Diation Effects Facility (RADEF) of the University of Jyvaskyla, Finland.The device was tested in a 50-MeV proton beam for total dose up to 10 krad,which it survived without observable degradation [39]. As the device relieson a commercial version of the Xilinx Virtex-4 field-programmable gate ar-ray, we have implemented a triple-redundant memory with active scrubbingrunning parallel to the normal operations of the instrument [43]. The sys-tem was tested in RADEF to be able to cope with a 50-MeV proton fluxof 10 cm − s − , after which the rate of double bit errors became significant[39].The instrument, being integrated into the satellite short stack, is alsosensitive to electromagnetic interference. Especially the scintillator detectorsignal path is affected by the electromagnetic emission of other spacecraftsubsystems. This has led to an increase of the noise levels in this signal andthe inability to detect at the smallest signal levels, which has increased thethreshold of the electron measurements from the nominal 0.7 MeV [39] to 1.5MeV achieved in space [40]. The plasma brake payload is based on the Coulomb drag principle, whichis the driving phenomenon behind the Electric Solar Wind Sail (E-sail) inven-tion [44, 45]. The brake itself consists of a 100 meter long tether; a storagereel; a vacuum qualified piezo motor and control electronics for tether de-ployment; a high voltage source; and four electron guns. Once the tether hasbeen deployed, it can be charged with a voltage of either +1 kV or − × ×
10 cm. The spacecraft EPS supplied 3.3 V, 5 V,and 12 V to EPB. Power consumption of EPB depends on the operationmode: launch locks use 1.25 W each for about 20 sec (locks are not releasedat the same time), high voltage tether system consumes a few hundreds ofmW depending on the ambient ionospheric plasma density, and the reeling14 igure 5: Tether Reel FM board from both sides. On the left, tether reel lock Kieku readyand locked. The black object left from Kieku is the optical feedback Kyyl¨a. The electronguns are located satellite side panel X+, next to the payload which deploys from Z- endof the satellite. system draws 2.3 W during the deployment. None of these tasks are executedsimultaneously, and the tether is not deployed all at once. The data rate islow throughout the mission as only the tether voltage and current are sampledwith a frequency of 10 Hz.The tether itself is constructed of four aluminum filaments and it is basedon the Heytether geometry [46]. The tether is deployed with the help ofcentrifugal force, the satellite must therefore be spinning around a suitableaxis. Once the proper spin mode is reached, the tether reel motor is activatedand the tether is slowly unreeled out to space. At the tip of the tether thereis an aluminum tip mass, whose task is to assist in tether deployment byincreasing the pull force experienced by the tether. On the bottom side of thereel there is a slip ring serving as the contact point for the high voltage sourcethrough two cantilever spring sliders being the only mechanically redundantsubsystem of EPB. When a positive voltage is applied, one or more electronguns are activated in order to eject excess electrons and thus maintain thepositive voltage, as the surrounding plasma attempts to neutralize it. Innegative tether voltage mode, the tether gathers positive ions from the plasmaand the conducting parts of the satellite surface collect the same flux ofthermal electrons from the plasma to maintain current balance..The plasma brake payload was tested prior to the system level tests. Vi-bration tests were carried out to qualify the mechanical components, PCBs,15nd the reel motor, especially, as the motor was designed for laboratory use.It was noted that the high voltage sliders dug two dents to the slip ring thatwere able to stop reel rotation. Simple resistor-based launch locks were in-troduced to the bottom side of the reel PCB to keep the sliders apart fromthe slip ring during the launch. The functionality of the payload was success-fully tested in thermal-vacuum. Furthermore, specific to EPB payload, highvoltage tests were made, and the tether outreeling was tested to determinethe minimum centrifugal force required for the tether deployment.
The miniaturized spectral imager, AaSI, as shown in Fig. 6, is the mainpayload of the Aalto-1 nanosatellite. The imager is based on a tunable FPI,which is used as an adjustable passband filter. This enables the imager toacquire images at freely selectable wavelengths. The operational range is500–900 nm and the spectral resolution is 10–20 nm. In addition to thespectral imager, a visible (VIS) spectrum Red–Green–Blue (RGB) camera isincluded in the instrument [47, 48, 49].
Table 1: Main parameters of AaSI.
Wavelength range
Spectral resolution
Field of view ◦ × ◦ (SPE), 15 ◦ × ◦ (VIS) Spectral image size ×
512 pixels
VIS image size × Number of spectral bands
6, 25 or 75
Size × ×
48 mm Mass
600 gTable 1 introduces the main parameters of AaSI.16 igure 6: The Aalto-1 Spectral Imager AaSI. The size of the instrument is ca. 0.5 U andit is compatible with the PC104 interface. The instrument has two cameras: a visiblespectrum RGB camera (left) and a spectral imager (right). . Platform The Aalto-1 platform subsystems include an EPS [50], an ADCS [51], aGPS-based navigation system [52], a UHF [53, 54] and S-band [55] radiosfor Telemetry, Telecommand & Communication (TT&C), and a Linux-basedOBC [56, 57, 58, 59]. The electronic subsystems are placed in two circuitboard stacks, the Long Stack and the Short Stack, which are connected usinga stack interface board. The electronics followed CubeSat electronics format,whereas the bus pin-layout followed PC-104 standard.The design philosophy of the CubeSat platform is a hybrid combination ofsubsystems developed in-house and commercial products. The satellite struc-ture, solar panels, Sun sensors, TT&C and OBC were fully designed in-housewhereas the ADCS and the EPS were procured from commercial partners.The CubeSat structure, antenna and antenna deployment system were alsodeveloped in house. The in-house developed subsystems were fully designed,integrated and testing by student teams. The PCB designs were manufac-tured by commercial PCB provider, whereas the component soldering andstuffing was performed in our facility. Special consideration was employedin the design of the critical subset of subsystems, consisting of EPS, OBCand UHF. Redundant parts, fault detection and recovery procedures wereadded to increase their reliability and fault tolerance. The agile developmentapproaches were followed in the design and verification of the satellite. Thedevelopment process of the subsystems has been iterative, since the proto-type of each subsystem was developed and qualified in quick iterations. Thewaterfall verification approach was followed in the Flatsat, EQM and FM in-tegration [60]. The detailed design description of each platform subsystem ispresented in the subsequent subsections. The in-orbit performance of majorplatform subsystems can be read from [27].
The EPS ensures the power generation, conditioning, storage and distri-bution to each subsystem and payload [61]. The EPS was procured froma commercial partner Clyde Space (Currently ˚AAC-Clyde). The solar pan-els were designed in-house in order to accommodate the conductive surfacerequirements of EPB. A block diagram of Aalto-1 EPS is presented in Fig. 7The incident solar radiation is converted to electrical power by the solarpanels, developed at Aalto University [62]. The Solar Panel design featuredthermally conductive PCB design. In-house made design provided freedom18 igure 7: Block diagram of EPS on sensor and payload location and satellite structure design. The powerfrom panels is transferred to the Electrical Power System Control Board(EPSCB) where Battery Charge Regulators (BCR) convert the input voltagefrom the solar panels to battery charging voltage (6.2 V to 8.4 V). The PowerConditioning Modules (PCMs) are responsible for regulating the voltagesproduced by solar panels and the battery unit. The dc-dc converters convertthe voltage levels to the ones used by subsystems and distribute the powerto the Satellite Bus (SB). The major subsystems have dedicated power lines,which are controlled by switches located on the battery board and accessiblethrough stack connector. The operating voltages, standby and peak powerconsumption of platform avionics and payloads are provided in Table 2.The EPS has several safety features implemented for increased reliabilityof the platform. It monitors the I C bus lines for inactivity and erroneousbehaviour (see Fig. 3), which if detected will cause a power cycle event of thewhole platform. Battery power level is monitored as well and the low powermode is activated if depth of discharge is below the critical value. In thismode only the EPS is active, operating the battery charging circuits. Lastly,a timer feature, which starts a 30 minutes countdown after the first EPSpower up, was set as a redundant antenna deployment trigger, in addition tothe main dedicated countdown timers.19 able 2: Aalto-1 Power budget.
System Details Operatingvoltage (V) StandbyPower (W) PeakPower (W)TT&C UHF 12 0.2 1.55S-band 3.3 0 3.5ADS 12 0 7GPS Active Antenna 0.03 0.03GPS Receiver 0.015 0.1OBC 0.25 0.55ADCS Coils andElectronics 5 0.5 1.8Sun sensors 5 0 0.06Payloads RADMON 12, 5 0 1EPB 12, 5, 3.3 2.3 3AaSI 12, 5 0 4Total 3.295 22.59
The ADCS is the most critical subsystem to ensure the required point-ing and spin modes for payloads. Aalto-1 ADCS (iADCS100), provided bycommercial partners Berlin Space Technologies (BST) and Hyperion Tech-nologies, consists of an integrated solution of attitude determination sensorsand attitude control actuators. The attitude sensors include, gyroscopes,magnetometers and a star tracker. The Sun sensors were developed in-houseby Aalto University and integrated to the solar panels[63]. The attitude ac-tuators include magnetorquer rods and reaction wheels.Aalto-1 was the firstsatellite carrying iADCS100 attitude system and the Aalto students partici-pated in the development. The FM of Aalto-1 ADCS is shown in Fig. 8.
The GPS subsystem of Aalto-1 is shown in Fig. 9 which contains a FastraxIT03 GPS receiver and an Adactus ADA-15S patch antenna. When operated,the GPS subsystem consumes approximately 160 mW of power [64, 52]. Themain purpose of the subsystem has been to provide more accurate positioningthan Two Line Element (TLE)-based solutions, for example, during plasmabrake operations. The Fastrax receiver was selected as the manufacturer20 igure 8: FM of the ADCS iADCS100 with star tracker by Berlin Space TechnologiesGmbH and Hyperion Technologies. was willing to provide the receiver without the usual altitude and velocityrestrictions [65]. In early 2010s, when a GPS subsystem was included inthe Aalto-1 design, there were not many GNSS subsystems for nanosatellitesavailable as commercial off-the-shelf products.
A UHF transceiver was used as the primary radio on the Aalto-1 satellite.The UHF radio supported transmission power of up to 1.2 W. The unit, asshown in Fig. 12, is fully redundant, equipped with two cold redundant TICC1125-based transceivers and an MSP430 microcontroller (MSP430FR5729).It is capable of half-duplex bidirectional communication at 437.220 MHz. TheUHF communication system is equipped with two dipole UHF antennas, eachconnected to one of the two redundant radios [53]. The OBC software andthe arbiter can perform the switching from active to redundant radio.A UHF antenna deployment system, as shown in Fig. 10, consists oftimer control board for antenna release and two L-shaped doors to keepthe antennas stowed during launch. After the spacecraft is deployed, theantenna release mechanism burns the dyneema strings thereby deployingthe antennas. Additionally, the redundant timer on the EPS can trigger theantenna deployment. The deployed antenna configuration is shown in Fig. 11.21 igure 9: Aalto-1 S-band transmitter and GPS subsytem
An automatic UHF beacon is transmitted every two minutes by default. TheUHF beacon containing a static Morse code is transmitted every two minutesby default.Along with the UHF radio, Aalto-1 has an S-Band transmitter used forhigh speed telemetry downlink. Because of regulations, the S-band trans-mission can be active only above the Aalto Ground Station. The S-bandtransmitter featuring a single transceiver (TI CC2500) and a microcontroller(MSP430FR5739) is shown in Fig. 9. The communication frequency is 2.402GHz with the design data rate of 500 kbps. The S-band communication sys-tem uses an in house designed single circular polarization patch antenna. Italso forms a secondary downlink channel [55].
The Aalto-1 OBC consists of two cold-redundant 32-bit AT91RM9200,microcontrollers from Mircochip. The architecture hosts a 256-Mbit Syn-chronous Dynamic Random Access Memory (SDRAM) volatile memory (AS4C16M16S),22 igure 10: Aalto-1 Antenna deployment system and UHF-band antenna in stowed config-uration a parallel/NOR flash (S29JL064J), a dataflash (AT45DB642D), and a NANDflash (S34ML02G1). These different memories are used to store boot-loaders,kernel images and file systems. The architecture uses three different bus inter-faces including I C, UART and SPI. The UART, SPI and USB are supportedby the processor itself, while I C is handled by an an external controller(PCA9665).The OBC consists of several components that can be classified as watch-dogs, the most important one being the arbiter [57]. An MSP430-basedarbiter selects which of the two processors is powered therefore preventingmission failure due to hard failure of one of the OBCs. In the arbitrationlogic, a full reboot is required to switch from the active to redundant OBC.The switching procedure was set to execute when the arbiter powers up andit does not receive a heartbeat signal of the active OBC. A further readon system description, arbitration logic, Failure mode and effects analysis(FMEA) and error handling procedures, can be found in [56]. A detailed23 igure 11: Aalto-1 UHF antenna in deployed configuration block diagram of the OBC representing the data interfaces with payloadsand platform subsystems is shown in Fig. 3 whereas the flight spare modelof the OBC is shown in Fig. 13.The OBC runs the Linux operating system and bash for scripting dif-ferent command sequences. At the time of its selection in 2010, Linux wasnot a common choice for satellite OBCs, but has since become popular insmall spacecraft [66], [59]. Software of the OBC, due to high complexity ofthe Linux operating system, has been thoroughly analysed and additionallystrengthened against various identified failure scenarios[57]. The version con-trol in software development approach was followed in the software designwith regular commits to the Github repository.
Several on-board data handling tools have been used in existing Cube-Sat designs [67]. The Aalto-1 on-board data handling and flight software isbuilt around applications running on Linux which was quite a new choice fornanosatellites at the design selection stage. The applications utilise certainlibraries to communicate with satellite subsystems and the satellite internaldata bus. 24 igure 12: Aalto-1 UHF-band cold redundant transceiver
A number of libraries were developed for several subsystems, the mostprominent being libarbiter for arbiter, libeps for communication with EPS,libicp for communications with subsystems on I C and libradio and libsbandfor radio communication. The detailed description on design choices andlessons learned on Aalto-1 software design approach can be followed in [59].Linux is a feature full operating system with mature, stable and time-proven core code base. This simplified the development of flight logic andutilities. Well known Linux library ecosystem and APIs assured a properseparation of concerns.Nonetheless, Linux is a complex software which necessitated a thoroughanalysis to ensure reliability on the OBC[57]. During the system’s boot pro-cedure there is no possibility for intervention and thus needed to be madefault tolerant by adding an emergency boot procedure with reduced func-tionality and less dependencies. A memory storage was divided into sectionswith primary and recovery file systems. Unsorted Block Image File System25 igure 13: Aalto-1 On-Board Computer’s flight spare (UBIFS) have been used and it supports wear leveling and due to its use ofjournals is power loss tolerant. On the overall system level, a number of soft-ware and hardware watchdog timers are used in conjunction with the arbiterheartbeat output. Bus and radio communication libraries are strengthenedwith appropriate checksum and implementing non-blocking procedures.
There are two long and a short standard PC-104 stack to route power anddata signals among platforms and payloads . The long stack, required by fewsubsystems, is 2U long whereas the short stack is 1U long. As evident fromFig.2, the orientation of subsystems on one unit is different than those onthe other two units, therefore a stack interface board was used. The in-housebuilt structure is compatible with standard dimensions and provides mechan-ical interface to internal subsystems, solar panels and antenna deployment.The breakdown of total spacecraft mass is provided in Table 326 able 3: Aalto-1 mass budget
System Details Quantity Unit mass (g) Total mass (g)Structure 3U and harness 1 1180 1180TT&C UHF antenna 4 1 4UHF transceiver 1 90 90S-band & GPS board 1 75 75S-band antenna 1 50 50EPS Solar panels 4 130 520Control board 1 83 83Batteries 1 258 258OBC 1 75 75ADCS Coils and Electronics 1 360 360Sun sensors 6 10 60Payloads RADMON 1 360 360EPB 1 300 300AaSI 1 600 600Total 3572 4015The spacecraft used passive thermal control system [68]. The structurerails were anodized black since it provides optimum emissivity/absorptivityratio. The electrically conductive surfaces were masked before anodizationand later chromate coated. The unused areas of solar cell PCBs were goldplated. In order to increase the thermal conductivity from solar cell to thestructure, indium foil washers were placed in the screw joints. For betterthermal conductivity, many grounding vias were also placed in the solarcell footprints. The telemetry data of Aalto 1 reveals that the equilibriumtemperature is well maintained inside the spacecraft.
7. Satellite integration & testing
The model philosophy of the project followed a Flatsat, EQM and FMapproach. This approach was selected mainly due to the fact that all sub-systems and payloads were new development items and early verification ofthem was seen as highly beneficial. Additionally, lessons learned from otherCubeSat projects in other universities often highlighted the importance ofleaving significant amount of time for the integration and testing campaignon the system level prior to a launch.27 igure 14: Aalto-1 Flatsat model integrating (starting from top left continuing clockwise)AaSI, EPB, UHF, S-band with patch antenna and GPS, OBC and commercial ADCS
Most of the testing performed prior to system level integration was doneon subsystem level by each development team. Usually, development kitsand other test equipment was utilized rather than other satellite subsystemsas their development was performed concurrently by separate teams.The firstfull interface tests were performed on the Flatsat model which is shown inFig. 14. A number of interface mismatches were identified and troubleshootedat this stage. While being rather typical for any project with many concur-rent developments, earlier testing of the system as a whole, if possible, wouldprobably have saved required redesign and manufacturing effort at the EQMlevel. As most of the satellite subsystems were in-house developed, makingsmall modifications was however relatively easy at this point of the project.A full environmental qualification test campaign was performed with thesatellite EQM. No major issues were found during these tests, which raisedthe confidence level on the system level design. However, not all satellitefunctionalities had been implemented at this point and thus were not fullyreference tested prior and during the environmental testing campaign. This28eft some uncertainties to be fully verified later at the FM test campaign. Italso highlighted the importance of thorough reference testing and reportingprior to environmental tests.The satellite FM was built soon after the EQM environmental test cam-paign, including some necessary minor modifications. Testing with the EQMalso continued throughout the FM campaign and allowed simpler softwaredevelopment and testing on the system level, as well as tests that could havecaused unnecessary stress to the FM. Such tests were, for example, longduration durability testing, outdoor long-range testing and magnetic testing.Some issues were still found using the EQM and it was possible to implementnecessary fixes to the FM. One of such issues was related to a component inthe Telemetry/Telecommand (TM/TC) radio and may not have been noticedwithout the long duration durability testing, and could possibly have causedmission failure soon after the launch.A full acceptance test campaign was performed with the satellite FM.The FM testing consisted of pre-built test scripts to command the subsys-tems and receive respective telemetries. No major issues were found duringthese tests, as was expected thanks to the successful EQM tests. Due tothe late readiness of the third party provided ADCS and flight software, itwas not possible to perform a thorough enough functional or performancetest campaign for it. Testing of the ADCS algorithms was planned to beperformed using a hardware-in-loop approach, which did not work as ex-pected through the satellite main communication bus due to communicationdelays. Rather, access to the ADCS internal sensor and actuator bus wouldhave been required, but it was not possible at that point. This highlightedthe importance of early delivery of third-party systems with final and fullytested flight software and should be considered a high risk regarding any newdevelopments by a third party.The Aalto-1 launch campaign started after the assembly, integration andverification stage. A lot of issues were addressed even when the satellite wasin the launch pod. As an example, the batteries had become empty and itwas a trouble charging them because no such interface was provided on theaccess port. The batteries in FM were charged with a solar lamp transportedto the launch pod.A photograph of the integration of the FM into the commercial orbitaldeployer is shown in Fig. 15.In the end, the selected model philosophy proved to be a very suitable ap-proach for the project that had significant amount of new development items.29 igure 15: Aalto-1 integration with the deployer
Like in many other CubeSat projects, schedule issues were encountered toperform system level testing in the most ideal and thorough way possible.Emphasis on system level testing from the very beginning of developmentcan solve some of such issues encountered at later stages of the project. Amost-viable-product approach, used typically in software engineering, hasbeen followed and determined beneficial in projects after Aalto-1, where thehighest importance functions of the satellite are implemented and tested asearly as possible on the system level, and later incremented with additionalfeatures in order of priority towards the full satellite integration and testing.Such an approach however requires agile in-house development and closecooperation with third parties. Ultimately, the most suitable developmentapproach for any CubeSat project is highly dependant on many aspects, suchas the available resources, experience, the number of development items andthe usage of in-house or third-party systems. The development approachshould be carefully planned only after such aspects have been identified.
8. Ground segment & services
The ground segment originally used an Icom IC-910H radio transceiverand a relay based pre-amplifier that was designed for voice communication.Reception was implemented using an RTL-SDR. Five months after launch,the setup was updated with a newly developed solid state switched pre-amplifier due to problems with the relay-switched pre-amplifier.30n 2018, the transceiver was changed to a USRP B200 Software DefinedRadio (SDR). To ease operation of multiple missions from the ground sta-tion, the digitized radio signal is distributed to multiple programs througha shared memory buffer. With the OpenWebRX software, the spectrumbetween 431 MHz and 439 MHz can be monitored using a web browser.Brushed motors in the antenna rotator caused strong, broadband inter-ference close to the antenna while rotating during a satellite pass. The issuewas reduced with upgraded rotators and an upgraded controller.Block diagram of the relevant parts of the currently operational groundstation is presented in Fig.16.
Figure 16: A block diagram of Ground station.
The ground segment is controlled using the Mission Control Center (MCC)software developed by the Aalto-1 team. The back-end is based on a Post-31reSQL database that stores every received packet with a timestamp. Fur-thermore, the housekeeping system stores every housekeeping value sepa-rately with timestamp which has proven cumbersome and ineffective due tostorage of large set of values in the database. For upcoming missions, we planto use a database structure that uses one table per subsystem and one row in-cludes an entire housekeeping package of that subsystem. This should reducequery time substantially as not every value has to be queried independently.We have developed a Graphical User Interface (GUI) for the MCC basedon Qt. Qt was chosen since the application programming interface (API) doesnot change very quick which ensures long compatibility in the future. TheGUI shows information about the current position of the satellite, satellitepasses, received and transmitted packets and housekeeping most recent data.In addition, the history of a single housekeeping value can be plotted.
9. Conclusion
The Aalto-1 projects key scientific and technological and educational ob-jectives were achieved. The platform and payloads were successfully de-signed, developed and integrated with many student teams getting hands-onlearning. The integration, testing, verification and launch activities were suc-cessfully accomplished. The subsystems and payloads demonstrated partialmission success with many lessons learned which have been briefed in anaccompanying paper.This project started a new era of space activities in Finland. A numberof new space start-ups were founded as an outcome of this project. The(former) Aalto satellites group members have started and joined a number ofnew missions, such as ICEYE SAR satellite constellation, Aalto-3, ReaktorHello World, FORESAIL [69], and Comet Interceptor [70]. The Aalto-1design has been been beneficial in the space technology curriculum and asource of inspiration for new students in the space technology lab.
Acknowledgements
The RADMON team thanks P.-O. Eriksson and S. Johansson at the Ac-celerator Laboratory, ˚Abo Akademi University, for operating the cyclotron.Testing work at the University of Jyvaskyla has been supported by theAcademy of Finland under the Finnish Centre of Excellence Programms32006-2011 and 2012-2017 (Project No:s 213503 and 2513553, Nuclear and Ac-celerator Based Physics), and by the European Space Agency (ESA/ESTECContract 18197/04/NL/CP).Aalto University and its Multidisciplinary Institute of Digitalisation andEnergy are thanked for Aalto-1 project funding, as are Aalto University,Nokia, SSF, the University of Turku and RUAG Space for supporting thelaunch of Aalto-1.
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