Innocube—A Wireless Satellite Platform to Demonstrate Innovative Technologies

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Article InnoCube—A Wireless Satellite Platform to Demonstrate Innovative Technologies

Benjamin Grzesik 1,2,* , Tom Baumann 3, Thomas Walter 3 , Frank Flederer 3, Felix Sittner 3, Erik Dilger 3, Simon Gläsner 1, Jan-Luca Kirchler 1, Marvyn Tedsen 1, Sergio Montenegro 3 and Enrico Stoll 2

1 Institute of Space Systems (IRAS), TU Braunschweig, 38108 Braunschweig, Germany; [email protected] (S.G.); [email protected] (J.-L.K.); [email protected] (M.T.) 2 Chair of Space Technology, TU Berlin, 10587 Berlin, Germany; [email protected] 3 Chair of Computer Science VIII, Aerospace Information Technology, University Wuerzburg, 97070 Würzburg, Germany; [email protected] (T.B.); [email protected] (T.W.); [email protected] (F.F.); [email protected] (F.S.); [email protected] (E.D.); [email protected] (S.M.) * Correspondence: [email protected]; Tel.: +49-(0)531-391-9988

Abstract: A new innovative satellite mission, the Innovative CubeSat for Education (InnoCube), is addressed. The goal of the mission is to demonstrate “the wireless satellite”, which replaces the data harness by robust, high-speed, real-time, very short-range radio communications using the SKITH (SKIpTheHarness) technology. This will make InnoCube the ﬁrst wireless satellite in history. Another technology demonstration is an experimental energy-storing satellite structure that was developed in the previous Wall#E project and might replace conventional battery technology in the Citation: Grzesik, B.; Baumann, T.; future. As a further payload, the hardware for the concept of a software-based solution for receiving Walter, T.; Flederer, F.; Sittner, F.; Dilger, E.; Gläsner, S.; Kirchler, J.-L.; signals from Global Navigation Satellite Systems (GNSS) will be developed to enable precise position Tedsen, M.; Montenegro, S.; et al. determination of the CubeSat. Aside from technical goals this work aims to be of use in the teaching of InnoCube—A Wireless Satellite engineering skills and practical sustainable education of students, important technical and scientiﬁc Platform to Demonstrate Innovative publications, and the increase of university skills. This article gives an overview of the overall design Technologies. Aerospace 2021, 8, 127. of the InnoCube. https://doi.org/10.3390/aerospace 8050127 Keywords: CubeSat; wireless-bus; harness free satellite; satellite technology; CubeSat GNSS; laser ranging; structural battery; dependable software; Rodos Academic Editor: Paolo Tortora

Received: 28 February 2021 Accepted: 29 April 2021 1. Introduction Published: 4 May 2021 CubeSats have evolved signiﬁcantly over the last 10 years. While they were once used

Publisher’s Note: MDPI stays neutral primarily for educational purposes, they are now capable of performing more complex with regard to jurisdictional claims in tasks. The goal of the research at the Institute of Space Systems of Technical University of published maps and institutional afﬁl- Braunschweig (IRAS) is to further enhance the performance of small satellites. In order to iations. tackle even more complex tasks, such as supporting active debris removal (ADR) or scien- tiﬁc missions, new technologies are implemented. The chair of Information Technology for Aerospace at the University of Wuerzburg is mainly researching reliable distributed and autonomous computing systems, focusing on space and aeronautic projects. Both universities partnered and are jointly developing a 3U–CubeSat for innovative technology Copyright: © 2021 by the authors. Licensee MDPI, Basel, Switzerland. advancement and practical, sustainable education of students. The aim of the project is the This article is an open access article technology demonstration of two innovative technologies, which have been previously distributed under the terms and developed and funded by German Aerospace Center (DLR) InnoSpace Masters program: conditions of the Creative Commons SKITH (SKIpTheHarness) and Wall#E. InnoCube concludes its three year development Attribution (CC BY) license (https:// phase with the launch into orbit in 2023, aiming for a mission of at least one year. The creativecommons.org/licenses/by/ planned CubeSat will be developed by students and research staff, going through all phases 4.0/). of a space project life cycle. The following article describes the mission framework with

Aerospace 2021, 8, 127. https://doi.org/10.3390/aerospace8050127 https://www.mdpi.com/journal/aerospace Aerospace 2021, 8, 127 2 of 20

a focus on the design criteria and operational concept. Therefore, Section2 covers the mission, payloads and boundary conditions, Section3 gives an overview of the satellite bus including the software, while Section4 gives a summary and outlook.

2. Mission The goals of the CubeSat project can be divided into four categories: • Long-term goals; • Technological goals; • Scientific goals; • Educational goals. The long-term goal is to increase both universities competencies and to raise TRL of used components for further development. This is reflected in the technology goals to demonstrate SKITH in an operational environment and have the first wireless satellite bus with respect to data harness. A further goal is the assessment of the functionality of a new battery concept of a carbon-reinforced plastic structural battery in a relevant space environment. In addition, a concept of an on-orbit navigation laboratory using a software-defined Global Navigation Satellite System (GNSS) receiver and Laser Ranging Reflector Experiment (LRR) is flown. The scientific goals comprise the validation of the initial theoretical assumption and to validate the GNSS experiment using the accuracy of a miniaturized laser retro reflector for CubeSats. Additionally, if possible the LRR measurements analyses should be used for improvement in positioning accuracy applying in orbit updates to the navigation algorithms. With the anticipated highly accurate orbit data, propagators developed at IRAS [1,2], can be improved. Works in the field of space situational awareness, rendezvous, and docking for active debris removal missions can be advanced. In addition to the scientific goals, one of the main objectives (educational goals) of the project is the training of students. This includes not only component development and technical applications, but also satellite operations. In summary, the design, fabrication, launch and operation of a small satellite with the following boundary conditions are to take place within 4 years: • Fabrication, validation and verification of a CubeSat flight model within a maximum of 3 years; • Successful launch with up to three months of Launch and Early Orbit Phase (LEOP) and payload calibration; • Mission life expectancy of 12 months (in-orbit); • Verify and validate science experiments and record science data through end of mission(active) and through end of life (passive–LRR); • Verify and validate technologies (SKITH and WALL#E); • Hardware and software testing; • Satellite operations with students.

2.1. Orbit Design The InnoCube satellite will not be launched with a dedicated launcher; instead, it will be brought into orbit as a secondary payload. A voluntary orbit cannot be chosen. Therefore, different launch opportunities are being examined. As the number of CubeSats launched have increased rapidly in the last few years, various vendors offer launch services for small satellites with different orbital parameters, perhaps the most common in popular media being SpaceX and its capabilities to offer piggyback services onboard their Starlink mission. The major requirements considering orbit design are defined by the mission lifetime. To fulfill a minimum on-orbit mission of one year, a minimum orbit altitude of 400 km is defined. A maximum altitude of 600 km is defined by compliance with the Space Debris Mitigation guidelines, adhering to deorbiting not later than 25 years after the end of operational life. Aerospace 2021, 8, x FOR PEER REVIEW 3 of 20

2.1.1. Orbit Analysis An analysis is performed concerning the orbital requirements for the payload, espe- Aerospace 2021, 8, 127 cially the laser ranging reflector and GNSS experiment, and the thermal environment 3of of 20 the satellite. The laser ranging experiment can be operated at an altitude ranging from 400 km to 600 km with no prerequisites regarding inclination. Thus, the limiting constraint is the 2.1.1.accessibility Orbit Analysis to the ground station in Braunschweig. AnThe analysis GNSS payload is performed can provide concerning continuous the orbital availability requirements from a large for the number payload, of re- espe- ciallyceivable the lasersatellites ranging due to reﬂector low interference and GNSS an experiment,d good reception and thegeometry thermal at environmentall times on a of thelow satellite. earth orbit below 3000 km ([3], see Figure 1 left). The L1 band shared by the Global PositioningThe laser System ranging (GPS) experiment and Galileo can can be operatedprovide a atlarge an altitudenumber of ranging receivable from satellites 400 km to 600in the km nominal with no carrier prerequisites frequency regarding due to is inclination.wide main lobe Thus, [4]. the limiting constraint is the accessibilityA minimum to the inclination ground station is necessary in Braunschweig. to access the ground station in Braunschweig, whichThe is GNSSat a latitude payload of 52.26° can provide North. Apart continuous from financial availability considerations, from a large a maximization number of re- ceivableof access satellites time is the due key to consideration. low interference and good reception geometry at all times on a low earthThe different orbit below thermal 3000 conditions km ([3], see arising Figure from1 left). different The orbits L1 band are negligible, shared by as the eclipse Global Positioningdurations are System very (GPS)similar and in Galileolow altitude can provide orbits. No a large subsystem number of of th receivablee satellite satellitesneeds to in theoperate nominal in specified carrier frequency temperature due ranges. to is wide Passive main methods lobe [4 are]. being implemented to keep the satellite in safe temperature and power levels.

Figure 1. left.: Global Navigation Satellite System (GNSS) in orbit reception geometry [4]; right.: Figure 1. left.: Global Navigation Satellite System (GNSS) in orbit reception geometry [4]; right.: Visibility of GNSS satellites in different orbital altitudes; top: Global Positioning System (GPS); VisibilityMiddle: Galileo; of GNSS Bottom: satellites combined in different [3]. orbital altitudes; top: Global Positioning System (GPS); Middle: Galileo; Bottom: combined [3]. 2.1.2. Space Debris Mitigation Guidelines A minimum inclination is necessary to access the ground station in Braunschweig, IRAS as an institute researching the subject of space debris, the Space Debris Mitiga- which is at a latitude of 52.26◦ North. Apart from financial considerations, a maximization tion guidelines are an important aspect in the design and operation of a satellite. The ap- of access time is the key consideration. plicable guidelines include deorbiting not later than 25 years after the end of operational The different thermal conditions arising from different orbits are negligible, as eclipse life, a re-entry burn up above 15 km, a minimization of risk of explosion and no generation durations are very similar in low altitude orbits. No subsystem of the satellite needs to of space debris. operate in specified temperature ranges. Passive methods are being implemented to keep The analysis tool DRAMA (Debris Risk Assessment and Mitigation Analyses) [5,6] the satellite in safe temperature and power levels. was used for the calculation of maximum lifetime after end of operations and to confirm 2.1.2.the maximum Space Debris possible Mitigation orbital alti Guidelinestude. The simulation resulted in a natural orbital lifetime of 17 years at an orbital altitude of 576 km (using the latest Two Line Elements (TLE) of the IRAShighest as Starlink an institute satellite researching [7]). Thus, the the subject space-debris of space debris,avoidance the guidelines Space Debris are Mitigation met by guidelinesa 5% margin. are anThe important simulated aspect individual in the compon designents and demise operation between of a satellite. an altitude The of applicable 75 km guidelinesand 80 km. include deorbiting not later than 25 years after the end of operational life, a re-entryThe burn current up abovesatellite 15 design km, a minimizationdoes neither include of risk ofany explosion deployables and nor no generationextendable of spacestructures, debris. except for the ultra-high frequency (UHF) antennas. In addition, the operation andThe evaluation analysis of tool the DRAMA scientific (Debrispayloads Risk can Assessment be expected and to Mitigationprovide highly Analyses) accurate [5 ,6] was used for the calculation of maximum lifetime after end of operations and to confirm the maximum possible orbital altitude. The simulation resulted in a natural orbital lifetime of 17 years at an orbital altitude of 576 km (using the latest Two Line Elements (TLE) of the

highest Starlink satellite [7]). Thus, the space-debris avoidance guidelines are met by a 5% margin. The simulated individual components demise between an altitude of 75 km and 80 km. The current satellite design does neither include any deployables nor extendable structures, except for the ultra-high frequency (UHF) antennas. In addition, the operation Aerospace 2021, 8, 127 4 of 20

and evaluation of the scientiﬁc payloads can be expected to provide highly accurate position data and reduce collision risk where appropriate. Thus, the requirements for minimizing the risk of explosion and no further creation of space debris are addressed by the satellite, mission design, and adherence to the CubeSat standard.

2.2. Communication Analysis To get knowledge about contact times between the ground station and the satellite, three main orbits that ﬁt in previous considerations are being analysed. • An orbit with an altitude of 550 km and an inclination of 53 degrees (SpaceX Rideshare /Starlink); • a sun synchronous orbit with an altitude of 550 km (various vendors, e.g., Vega Small Spacecraft Mission Service (SSMS), SpaceFlight Industries); • the orbit of the International Space Station (ISS) with an altitude of around 420 km and an inclination of 51.6 degrees (ISS NanoRacks CubeSat deployer). The nature of these low Earth orbits (LEO) indicates common similarities. All orbits feature ﬁve to six contact windows per day, spanning between seven and nine minutes. Sun synchronous orbits feature two contact windows per day, consisting of two to three orbits. Considering a maximization of communication time, the favoured orbit is the Starlink orbit, averaging 9.1 min per pass with a contact time per day of 53 min. Secondly, the sun synchronous orbit reaches an average 8 min per pass and 45 min per day, respectively. The lower ISS orbit reaches 7.3 min per pass and 38 min per day. With a proposed data rate of 9600 bps, data amounts of 2.5 to 3.4 megabytes per day are expected to be transferred.

Link Budget Analysis The quality of the proposed communication link is estimated with the formulation of conservative parameters, regarding all detrimental inﬂuences. The mission requires the communication system to perform at an elevation angle of 5◦ with an orbit altitude of 600 km. The system uses UHF frequencies of the amateur–radio band with a data rate of 9600 bps. The quality of the signal is parameterized with a bit error rate of less than 10−5. The ground station consists of a high gain antenna and a 120 W transmitter. Similar to most CubeSats, the satellite transmits with 1 W of power. The antenna of the satellite is a ground-plane antenna consisting of four monopoles, resulting in an omnidirectional radiation pattern with right–handed circular polarization. A strong signal can be achieved in the uplink path due to high transmitting power of the ground station. The calculated link margin is 26 dB. The downlink path is a lot more fragile, considering a noisy environment at the ground station and far less transmitting power at the satellite. Without polarization loss and the proposed data rate, a link margin of 5 dB is expected. The resulting margin in operation is expected to be higher, as the noise level of the ground station and the discrimination of the signal on the satellite need to be determined.

2.3. Power Budget Analysis The power that is generated by four body-mounted solar panels has been determined for multiple mission scenarios using the power simulation tool from the Virtual Satellite (VIRSAT) software [8]. These panels cover the whole surface of the CubeSat except the rails and standoffs According to typical solar panels [9,10] and the dimensions of the satellite we assume an area of 0.021 m2 with solar cells and an effectivity of about 29%. The power budget has been calculated for a 550 km sun-synchronous orbit in order to cover a wide spectrum of the possible launch opportunities for InnoCube. In this orbit the time of sunlight is 3585 s, the time of eclipse is 2120 s according to a Systems Tool Kit (STK) analysis. The worst case 24 h operational scenario is as follows: Aerospace 2021, 8, 127 5 of 20 Aerospace 2021, 8, x FOR PEER REVIEW 5 of 20

•• ElectricalElectrical Power Power System System (EPS), (EPS), Receiver Receiver (Rx), (Rx), Beacon Beacon und onboard und onboard computer computer (OBC) (OBC) alwaysalways ON; ON; •• TransmitterTransmitter (Tx) (Tx) 50 50min min within within 24 h 24 for h 10 for min 10 minON and ON battery and battery heater heater 75 min 75 ON; min ON; •• attitudeattitude determination determination and and control control system system (ADCS) (ADCS) low power low power Mode ModeON + 10 ON min + 10 min highhigh power power mode mode and and 120 120 min min medium medium power power mode; mode; •• GNSSGNSS Payload Payload 150 150 min min in ON in ON and andWall#E Wall#E charging charging in max in power max power mode modefor 15 h for per 15 h per 2424 h; h; •• OperationalOperational average average temperature temperature of the of solar the solar panel panel 50 °C. 50 ◦C. TheThe simulations simulations and and calculations calculations of the of the power power budget budget shows shows (Figure (Figure 2) that2) that the thesys- system temis power is power positive positive and and the the batteries batteries have have enough enough margin margin and and can can always always be be fully fully charged. charged.The average The average generated generated power power with the with assumptions the assumptions above above is about is about 67 Wh 67 in Wh 24 in h while24 the hsatellite while the needs satellite 54 Wh needs in the 54 describedWh in the worstdescribed case worst scenario. case The scenario. maximum The maximum power needed is powerabout needed 15 Wh is while about in 15 a speciﬁcWh while ADCS in a mode.specific Therefore,ADCS mode. the Therefore, average depth the average of discharge depth(DoD) of for discharge a battery (DoD) with for about a battery 77 Wh with capacity about [ 1177] Wh is <5% capacity with the[11] maximumis <5% with DoD the < 10%. maximumThis should DoD have <10%. a positive This should effect have on thea positive battery effect degradation on the battery and lifetime. degradation and lifetime.

FigureFigure 2. 2.WorstWorst case case power power simulation. simulation.

2.4.2.4. Scientific Scientific and and Technical Technical Goals Goals and and Payloads Payloads 2.4.1.2.4.1. SKITH SKITH (SKIpTheHarness) (SKIpTheHarness) SKITHSKITH is isthe the acronym acronym for forSKIpTheHarness, SKIpTheHarness, which which is a very is a compact very compact description description of of thethe main main idea. idea. Within Within the the InnoCube InnoCube satellite, satellite, a wireless a wireless data exchange data exchange between between all satellite all satellite modules,modules, i.e., i.e., the the onboard onboard computer computer (OBC), (OBC), ADCS, ADCS, power power distributi distributionon and conditioning and conditioning unitunit (PCDU), (PCDU), up-/downlink up-/downlink transceiver transceiver and the and payload the payload is realized. is realized. It is the It first is the time first a time a satellitesatellite has has been been designed designed following following this thisappr approach.oach. In this In mission this mission we aim we to aimdemonstrate to demonstrate thethe SKITH SKITH concept concept for for CubeSats. CubeSats. It features It features some some fundamental fundamental advantages advantages over a overcabled a cabled harness,harness, which which will will be beelaborated elaborated in the in thefollowing. following. TheThe communication communication between between on-board on-board componen componentsts within within conventional conventional satellites satellites is is basedbased on on a awired wired harness. harness. Standard Standard concepts concepts used used are point are point to point to pointconnections connections or bus or bus structuresstructures e.g., e.g., Controller Controller Area Area Network Network (CAN). (CAN). Known Known issues issues with cabled with cabled connections connections areare the the risk risk of cable of cable failures failures (due (dueto broken to broken wires or wires connectors), or connectors), a high mass a high (up to mass 10% (up to of10% total of satellite total satellite [12]) and [12 inflexibility]) and inflexibility to later design to later changes. design The changes. testability The of the testability harness of the on an integrated satellite requires additional hardware or software effort, to inspect all harness on an integrated satellite requires additional hardware or software effort, to inspect connections. all connections. To overcome these issues, the chair for aerospace information technology developed To overcome these issues, the chair for aerospace information technology developed a a concept that replaces all internal wired data connection by low power wireless interfaces.concept Subsequently, that replaces all all the internal aforementioned wired data failure connection risks byare low obsolete. power The wireless satellite interfaces. Subsequently, all the aforementioned failure risks are obsolete. The satellite integration becomes easier and errors due to inaccurately assembled connections are impossible, at least for data interfaces, as there is still the need for power transmission lines.

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Another advantage of a wireless communication concept especially in satellites is the simpliﬁed implementation of redundancy. The wireless concept shares some characteristics with a wired bus structure. The common media is not the wire, but the spectral range of the transmission frequency. In a wired bus, the risk of one broken node, blocking the whole bus by shorting the data lines to ground, needs to be handled. A common solution for this is duplicating the bus and dedicated modules. In a wireless concept the risk of blocking the communication channel by one broken module is not given, once this module is powered off. Thus, the only need is the ability to separately power off each module, a need that is common on satellites. Finally testing of the interfaces can be done without a physical access to the integrated satellite. It is as simple as placing a receiver in the surroundings and analyze the data exchange in the satellite. The SKITH interface on InnoCube will be realized by using an EFR32FG12 microcontroller by Silicon Labs. It is a commercial off the shelf (COTS) component that already includes a 2.4 GHz front end. All external elements needed are an antenna and a tuning circuit. Since the antenna is implemented on a circuit board (PCB) layout, there is only a small inductance that needs to be mounted aside the microcontroller. The HF modulation supported by the front-end is a Gaussian frequency shift keying (GFSK) or quaternary phase shift keying (QPSK). The maximum transmission power is 19 dBm, the receiver sensitivity for a GFSK modulation is −94.8 dBm at a data rate of 1 Mbps [13]. The board layout for a complete SKITH interface, including antenna, needs 25 mm by 30 mm and will be included on all modules. An image showing a PCB carrying two SKITH interfaces is shown in Section 3.5.1.

2.4.2. Radiation Test As the EFR32FG12 microcontroller was never used in space before it has been pre- liminarily radiation tested. Four off-the-shelf development boards of the microcontroller were irradiated by a Cobalt-60 gamma-radiation source. They were exposed to a total dose of 13.12 krad over 20 h. While testing the controllers were powered and exchanged radio messages between each other and performed regular integrity checks of all memories. The test results show no memory errors or communication failures. In a subsequent test another four controllers were to be irradiated until failure, but the test had to be aborted after 3.75 h due to time constraints. At the end, a total dose of 33.75 krad has been reached. Again, no errors in memory were observed. After a dose of 20 krad one controller showed less than 1% of radio packet loss, the others none. Reference [14] estimates 1.2 krad/year for 1mm shielding in LEO. With these test results the chip is expected to be suitable for at least 1 year planned mission lifetime with a good safety margin. Due to budget constraints, high-energy proton and ion tests are not planned, but these events are typically less of a concern in a short time LEO mission.

2.4.3. Redundancy Concept For the radio communication to work, it is important to always have only one master node active, that controls the communication. To ensure redundancy two master nodes are present but one active one is determined by the hardware switches in the power unit. The two master nodes are connected to the power by so-called keep-off switches. These switches keep their off-state if they receive regular pulses on their control input. If these pulses are lacking for a certain period they turn to the on state. Now one master node controls the switch of the other and vice versa. The active master node sends keep-off pulses to the other node’s switch, when it determines itself to be healthy and working. This ensures only one active master at a time. The generation of the pulse is placed in a piece of code that only executes when the radio communication is reliably working. This is the code that processes received frames. Other nodes may only send frames when they receive a sync frame from the master, so if the Aerospace 2021, 8, x FOR PEER REVIEW 7 of 20

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communication is reliably working. This is the code that processes received frames. Other nodes may only send frames when they receive a sync frame from the master, so if the mastermaster receivesreceives aa frameframe fromfrom anotheranother node,node, thisthis showsshows reliablereliable operationoperation ofof thethemaster’s master’s sendersender andand receiver.receiver. TheThe ﬂowflow ofof thethe signalssignals isis shownshown inin FigureFigure3 .3.

FigureFigure 3.3. OverviewOverview ofof thethe redundancyredundancy management.management.

IfIf anythinganything fails,fails, thethe keep-offkeep-off pulsespulses stopstop andand thethe otherother mastermaster nodenode bootsboots afterafter thethe timeout.timeout. IfIf thisthis nodenode isis working working it it will will itself itself send send a a keep-off keep-off to to the the ﬁrst first node node and and disable disable it. it. On power-on both nodes will boot at the same time. To ensure only one node stays on both masterOn power-on nodes areboth programmed nodes will toboot send at athe keep same off time. command To ensure at start-up only withone node a random stays delay.on both This master selects nodes one node are programmed randomly to beto activesend anda keep ensures off command no node is at preferred start-up inwith case a itrandom has unforeseen delay. This problems selects thatone causenode thisrandomly mechanism to be somehowactive and to ensures fail. no node is pre- ferredThe in activecase it masterhas unforeseen node manages problems the that redundancy cause this ofmechanism other nodes somehow in the satellite to fail. in software by listening for status messages sent over radio. As the master node is responsible The active master node manages the redundancy of other nodes in the satellite in for controlling the power unit it is able to switch all other nodes on and off as needed. software by listening for status messages sent over radio. As the master node is responsi- 2.4.4.ble for Radio controlling Communication the power Protocolunit it is able to switch all other nodes on and off as needed.

2.4.4.The Radio integrated Communication radio hardware Protocol in the EFR32FG12 microcontroller handles Open Systems Interconnection model (OSI) Layer 1 by itself. Further the hardware handles framingThe andintegrated checksum radio of hardware Layer 2. in The the software, EFR32FG12 therefore, microcontroller must handle handles radio Open packet Sys- collisiontems Interconnection avoidance and model addressing. (OSI) Layer 1 by itself. Further the hardware handles framing and Tochecksum keep things of Layer simple 2. andThe deterministicsoftware, ther thereefore, is onemust designated handle radio master packet node, collision which controlsavoidance all and communication addressing. as well as the power unit. To avoid that this master is a single pointTo of failurekeep things it is dual simple modular and deterministic redundant. Thethere master is one sends designated synchronization master node, frames which in regularcontrols intervals. all communication The structure as well of the as framesthe power send unit. on the To radioavoid link that is this outlined master in is Figure a single4. Thepoint sync of failure frames it contain is dual a modular list of relative redundan timest. toThe the master synchronization sends synchronization frame. These frames mark thein regular time slots intervals. at which The each structure other of node the frames may transmit. send on the The radio end oflink the is transmissionoutlined in Figure slot of4. The one sync node frames is marked contain by a the list beginning of relative oftimes the to next. the Thesynchronization slot lengthcan frame. be adjustedThese mark to accountthe time forslots different at which bandwidth each other requirements node may transmit. of the nodes. The end After of the the transmission slot-list the master slot of sendsone node its payload, is marked if any,by the like beginning any other of node. the ne Thisxt. canThe ensure slot length a minimum can be bandwidthadjusted to forac- everycount node.for different The maximum bandwidth latency requirements is depended of onthe the nodes. maximum After lengththe slot-list of the the cycle. master The framesends ofits thepayload, non-master if any, nodes like any is the other same node. except This that can the ensure time-slot a minimum list is missing. bandwidth for every node. The maximum latency is depended on the maximum length of the cycle. The frame of the non-master nodes is the same except that the time-slot list is missing. The payload contains the data of the higher level Rodos (real-time on-board dependable operating system) middleware protocol, which transmits the published topic data to all nodes. The Rodos middleware frames which contain the topic data are transmitted in the payload. One radio frame can contain multiple Rodos middleware frames to allow full usage of the slot if there are only small frames available. Addressing is then handled inside the Rodos middleware frames with its publisher/subscriber architecture and topic IDs. Each node processes all frames and forwards the topics it is subscribed to, to its local subscribers. The middleware and topics are further described in Section 3.5.5.

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FigureFigure 4. 4.Structure Structure of of implemented implemented radio radio protocol protocol frame. frame.

2.4.5.The WallE-2-Space payload contains the data of the higher level Rodos (real-time on-board dependable operating system) middleware protocol, which transmits the published topic data to The goal of the experiential payload WallE-2-Space is to examine previous researched all nodes. Wall#EThe technology Rodos middleware [15,16] in framesa relevant which space contain environment. the topic The data technology are transmitted integrates in the en- payload.ergy storage One functions radio frame into canthe containsupporting multiple structures Rodos of middlewarea spacecraft, equipping frames to allowthe carbon full usagefiber-reinforced of the slot ifplastics there are (CFRP) only small with frames solid-s available.tate battery Addressing materials is at then nanoscale handled and inside mi- thecroscale Rodos levels. middleware By partially frames replacing with its matrix publisher/subscriber polymer with active architecture material and for topicthe anode IDs. Eachand cathode, node processes solid electrolyte all frames and and conductive forwards theadditive, topics a it storage is subscribed function to, and to itselectron local subscribers.transportFigure 4. Structure can The be middleware achieved. of implemente In and orderd radio topics to protocol electrically are further frame. describedseparate the in Sectionanode and3.5.5 cathode. from each other, a separation layer made of a mechanically highly stressable glass fiber lies 2.4.5.between2.4.5. WallE-2-Space WallE-2-Space these two layers of carbon fiber laminate (Figure 5). TheThe goal goal of of the the experiential experiential payload payload WallE-2-Space WallE-2-Space is is to to examine examine previous previous researched researched Wall#EWall#E technology technology [ 15[15,16],16] inin aa relevant space space environment. environment. The The technology technology integrates integrates en- energyergy storage storage functions functions into into the the supporting supporting struct structuresures of a spacecraft, of a spacecraft, equipping equipping the carbon the carbonfiber-reinforced fiber-reinforced plastics plastics (CFRP) (CFRP) with withsolid-s solid-statetate battery battery materials materials at nanoscale at nanoscale and and mi- microscalecroscale levels. levels. By By partially partially replacing replacing matrix matrix polymer withwith activeactive materialmaterial for for the the anode anode andand cathode, cathode, solid solid electrolyte electrolyte and and conductive conductive additive, additive, a a storage storage function function and and electron electron transporttransport can can be be achieved. achieved. In In order order to to electrically electrically separate separate the the anode anode and and cathode cathode from from eacheach other, other, a a separation separation layer layer made made of of a a mechanically mechanically highly highly stressable stressable glass glass fiber fiber lies lies betweenbetween these these two two layers layers of of carbon carbon fiber fiber laminate laminate (Figure (Figure5). 5).

Figure 5. Schematic structure of an energy-storing fiber composite structure (Wall#E) [13].

Different experimental battery demonstrators are used for the satellite payload. The basis are CFRP structural batteries like that described above, with slightly varying layers based on the four types investigated in the original project. The layer structure of the Type I is based on the carbon fiber-based graphite-containing anodes, carbon fiber-based cathodes, and a commercial polyethylene oxide (PEO)-based separator. Type II consists of a layer structure of a carbon fiber-based cathode, the commercial separator and a lithium −2 metal anode. Type III uses a glass fiber-based separator with a basis weight of 105 g m impregnated with an electrolyte solution of PEO, lithiumbis(trifluormethylsulfonyl)imide Figure 5. Schematic structure of an energy-storing fiber composite structure (Wall#E) [13]. Figure(LiTFSI), 5. Schematic and acetonitrile. structure The of an same energy-storing functional ﬁber layers composite are used structure for type (Wall#E) IV but [13 with]. a basis Different experimental battery demonstrators are used for the satellite payload. The basis are CFRP structural batteries like that described above, with slightly varying layers based on the four types investigated in the original project. The layer structure of the Type I is based on the carbon fiber-based graphite-containing anodes, carbon fiber-based cathodes, and a commercial polyethylene oxide (PEO)-based separator. Type II consists of a layer structure of a carbon fiber-based cathode, the commercial separator and a lithium metal anode. Type III uses a glass fiber-based separator with a basis weight of 105 g m−2 impregnated with an electrolyte solution of PEO, lithiumbis(trifluormethylsulfonyl)imide (LiTFSI), and acetonitrile. The same functional layers are used for type IV but with a basis

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Different experimental battery demonstrators are used for the satellite payload. The basis are CFRP structural batteries like that described above, with slightly varying layers based on the four types investigated in the original project. The layer structure of the Type I is based on the carbon fiber-based graphite-containing anodes, carbon fiber-based cathodes, and a commercial polyethylene oxide (PEO)-based separator. Type II consists of a layer structure of a carbon fiber-based cathode, the commercial separator and a lithium metal anode. Type III uses a glass fiber-based separator with a basis weight of 105 g m−2 impregnated with an electrolyte solution of PEO, lithiumbis(trifluormethylsulfonyl)imide (LiTFSI), and acetonitrile. The same functional layers are used for type IV but with a basis weight of 25 g m−2 for the separator to improve the performance. The CFRP structure batteries will be mounted on parts of the satellite structure between wall and solar panels as well as inside the CubeSat, with some reference cells in pouch design. The WallE-2-Space payload will act as a nominal consumer to the satellite power bus, with the payloads own power distribution and charge regulator used to cyclize the battery and record the experimental data.

2.4.6. Experiment for Precise Orbit Determination (EPISODE) Experiment for Precise Orbit Determination (EPISODE) is a payload consisting of a GNSS receiver and miniaturized laser retro-reﬂector (LRR). The development is done through coursework and student work. The corresponding hardware for the GNSS receiver and the retro-reﬂectors is custom developed with COTS components. EPISODEs insights will support Space Debris research at the Institute of Space Systems. It further enables the provision of precise position data from the InnoCube, which contributes to a safer space environment by allowing collision warnings to be more accurately evaluated and assessed. The GNSS system consists of a front-end receiver with antennas for reception of GNSS satellite signals and a Field Programmable Gate Array (FPGA) based processing unit for generating the pseudo-random code, within the GPS system called the C/A (coarse acquisition) code, and correlation with the input signals. EPISODE will be composed of low-cost commercial off the shelf components and shall provide tracking solutions and position measurement data to the ground station at regular intervals. The main objective of the payload design is to provide a compact and low power solution for small spacecraft. For compactness, better power routing and less signal loss, a 4-layer circuit board design was chosen. As the antenna is one of the biggest parts of the receiving front-end, it will be placed on the outer structure of the spacecraft, instead of being integrated on the board. This leads to a smaller board and greater signal strength. The front-end utilizes a GNSS chip, designed for both industrial purposes with high requirements as well as for broad end-users, with an entire receiver chain on the chip. Its frequency range covers the L1/E1, B1 and G1 bands for GPS, Galileo, BeiDou and GLONASS. Currently, the combined system of front-end and processing unit is designed. For initial code development and testing, an Artix 7 FPGA with a Nexys 4 development board is used. The front-end receives a GNSS signal as C/A- code, where each transmitter has a unique code. An identical code must be generated and overlaid with the received code to correlate the relating satellite and determine its pseudo-range. To enable a position determination, four signals need to be correlated. Figure6 shows the necessary steps to determine the position from the acquired signal. To determine the position of the receiver, a triangulation problem has to be solved by the processing unit. The generation of the C/A code for GPS, using the developer FPGA, and matching to the received message was accomplished lately. Further developing steps include decoding of the navigation message and calculating the position data. The signal acquisition from Galileo and GLONASS satellites is implemented to compare different navigation systems. Aerospace 2021, 8, x FOR PEER REVIEW 10 of 20

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Figure 6. Steps that are necessary to achieve a navigation solution and tasks performed by an FPGA. FigureFigure 6. 6.Steps Steps that that are are necessary necessary to to achieve achieve a a navigation navigation solution solution and and tasks tasks performed performed by by an an FPGA. FPGA.As an alternative, the highly accurate, independent and supportive method of laser rangingAs an (SLR) alternative, will be used the highly for orbit accurate, determination. independent Its main and limitation supportive is methodpoor spatial of laser and rangingtemporalAs (SLR)an coverage, alternative, will be but used the SLR forhighly measurements orbit accurate, determination. independentare unambiguous Its main and limitation supportive and directly is poor method related spatial of to andlaser the temporalterrestrialranging (SLR) coverage, reference will be but system. used SLR for measurementsThis orbit knowledg determination. aree allows unambiguous Its independent main limitation and calibration directly is poor related andspatial valida- to theand terrestrialtiontemporal of the referencecoverage, on-board system. but GNSS SLR Thishardware. measurements knowledge A fund allows aramentale unambiguous independent requirement and calibration directlyfor a laser and related ranging validation to there- offlectorterrestrial the on-board is to reference provide GNSS system.observation hardware. This A throughknowledg fundamental thee allowsglobal requirement SLRindependent station for anetwork calibration laser ranging with and both reflector valida- high isaccuracytion to provideof the and on-board observation sufficient GNSS signal through hardware. strength. the global A Improvements fund SLRamental station inrequirement network laser ranging with for both measurement a laser high ranging accuracy tech- re- andnologyflector sufficient ismake to provide it signal appear strength.observation promising Improvements through to miniaturize the global in laserthe SLRlaser ranging station retro measurement reflector network of with a suitable technology both high size makeforaccuracy a CubeSat, it appear and sufficient enabling promising signal centimeter to miniaturizestrength. level Improvements measurements. the laser retro in reflectorlaserThe smaller ranging of a the measurement suitable number size of for tech-indi- a CubeSat,vidualnology prismsmake enabling it within appear centimeter such promising a reflector, level to measurements. miniaturize the higher the The achievablelaser smaller retro the reflectoraccuracy. number of An ofa suitableideal individual reflec- size prismsfortor awould, CubeSat, within therefore, such enabling a reflector, consist centimeter theof only higher level one the measurements. in achievabledividual prism. accuracy. The However, smaller An ideal the due reflectornumber to the would,oflimited indi- therefore,vidualfield of prisms view, consist thiswithin of solution only such one ais reflector, individualnot possible the prism. hiforgher low-flying However, the achievable satellites. due to accuracy. the A limited potential An field ideal solution of reflec- view, is thistora scaled solutionwould, reflector therefore, is not to possible CubeSat consist for dimensions,of low-flying only one satellites. similarindividual to A the potentialprism. Champ-Grace However, solution isLRR, due a scaled todeveloped the reflector limited by toGermanfield CubeSat of view, Research dimensions, this solutionCentre similar for is notGeosciences to thepossible Champ-Grace (GFZ),for low-flying consisting LRR, developedsatellites. of four A byprisms potential German shown solution Research in Fig- is Centre for Geosciences (GFZ), consisting of four prisms shown in Figure7. urea scaled 7. reflector to CubeSat dimensions, similar to the Champ-Grace LRR, developed by German Research Centre for Geosciences (GFZ), consisting of four prisms shown in Fig- ure 7.

Figure 7. left: Cube Laser Ranging Reflector Experiment (LRR) exemplarily positioned on CubeSat Figure 7. left: Cube Laser Ranging Reﬂector Experiment (LRR) exemplarily positioned on CubeSat top side; right: 3D printed CubeLRR reduced by approx. factor 4 with respect to the original sized top side; right: 3D printed CubeLRR reduced by approx. factor 4 with respect to the original sized (approx.Figure 7. 10left: cm Cube × 10 Lasercm × 10Ranging cm) Champ Reflector reflector Experiment holder. (LRR) exemplarily positioned on CubeSat (approx.top side; 10 right: cm ×3D10 printed cm × 10 CubeLRR cm) Champ reduced reﬂector by appr holder.ox. factor 4 with respect to the original sized (approx. 10 cm × 10 cm × 10 cm) Champ reflector holder.

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3. Spacecraft Overview The satellites avionics (see Figure 8) is composed of modules, that are common to most satellites. These include an onboard computer (OBC), an ADCS, a communication interface for up-/downlink and a PCDU. Most of the modules are dual modular redundant (DMR). The ADCS is replicated twice. Aside from the common SKITH interface implementing the control algorithms, there are additional components e.g., two micro electro mechanical systems (MEMS) gyroscopes, two magnetometer and front-ends for torquer control. Complementary, an interface to the sun-sensors and the reaction wheels is integrated. Aerospace 2021, 8, 127 Both ADCS modules are assembled on a single board, which is mounted in a self-con-11 of 20 tained unit, holding the torquer and wheels. Further details are given in the corresponding subsection of the ADCS system. The up-/down link transceivers for communication with the ground station are rep- 3. Spacecraft Overview licated to avoid a single point of failure. Additionally, a satellite modem is implemented, buildingThe satellitesa backup avionics communication (see Figure channel.8) is composed The modular of modules, concept that is not are commonlimited to to the most avi- satellites.onics system, These complementing include an onboard a modular computer structure, (OBC), anwhich ADCS, is explained a communication in the following interface forchapter. up-/downlink and a PCDU. Most of the modules are dual modular redundant (DMR).

FigureFigure 8.8.InnoCube InnoCube avionicsavionics overview.overview.

3.1. ModularThe ADCS Structure is replicated twice. Aside from the common SKITH interface implementing the controlInnoCube algorithms, is a 3U+ there CubeSat are additional(354 × 113 components× 113 mm³) with e.g., twoa mass micro not electroexceeding mechan- 4 kg. icalOne systems of the benefits (MEMS) of gyroscopes, a wireless bus two system magnetometer is the lack and of the front-ends data harness for torquer and the control. ability Complementary,to test hardware an in interface the laboratory to the sun-sensors as well as integrated and the reaction in the wheels satellite. is integrated.To utilize Boththose ADCSbenefits, modules the structure are assembled is designed on a singleto allow board, quick which access is and mounted exchange in a self-containedof subsystem mod- unit, holdingules, allowing the torquer for variable and wheels. circuit Further board detailssize, enabling are given COTS in the and corresponding custom designs, subsection which ofprovide the ADCS a certain system. radiation protection for the wireless bus modules. TheThe up-/downstructure concept link transceivers is to install for the communication subsystem circuit with boards the ground individually station arefrom repli- the catedfront towith avoid little a singleassembly point effort. of failure. For this Additionally, purpose, boards a satellite can be modem inserted is implemented,from the front buildingon evenly a spaced backup rails communication in the sidewalls channel. with clamping The modular mechanisms, concept wedgelocks, is not limited retaining to the avionicseach board system, in its complementinginserted rail slot aas modular shown in structure, Figure 9. whichWedgelocks is explained are pin-shaped in the follow- com- ingponents chapter. that expand when tightened, building a resistance against the side of the rail and board. This way, the circuit boards are secured in place providing resistance to shock and 3.1. Modular Structure vibration, as well as enabling a thermal transfer to the main structure. The power supply × × 3 to theInnoCube modules isis arealized 3U+ CubeSat with a backplane. (354 113 113 mm ) with a mass not exceeding 4 kg. One of the beneﬁts of a wireless bus system is the lack of the data harness and the ability to test hardware in the laboratory as well as integrated in the satellite. To utilize those beneﬁts, the structure is designed to allow quick access and exchange of subsystem modules, allowing for variable circuit board size, enabling COTS and custom designs, which provide a certain radiation protection for the wireless bus modules. The structure concept is to install the subsystem circuit boards individually from the front with little assembly effort. For this purpose, boards can be inserted from the front on evenly spaced rails in the sidewalls with clamping mechanisms, wedgelocks, retaining each board in its inserted rail slot as shown in Figure9. Wedgelocks are pin-shaped components that expand when tightened, building a resistance against the side of the rail and board. This way, the circuit boards are secured in place providing resistance to shock and vibration, as well as enabling a thermal transfer to the main structure. The power supply to the modules is realized with a backplane. Aerospace 2021, 8, 127 12 of 20 Aerospace 2021, 8, x FOR PEER REVIEW 12 of 20

Figure 9. Modular structure showing subsystem integration. Figure 9. Modular structure showing subsystem integration.

3.2.3.2. Thermal-System Thermal-System InIn general, general, InnoCube’s InnoCube’s thermal thermal control control design design has has to keep to keep the satellite’s the satellite’s internal internal tem- peraturetemperature at least at leastwithin within the range the rangeof −40 ofto −+8540 °C. to This +85 ◦rangeC. This derives range from derives the extensive from the useextensive of industrial use of grade industrial COTS grade components COTS components for the internal for the electronics. internal electronics.External units External (such asunits the (suchsun-sensors) as the sun-sensors)may be subjected may to be much subjected more toextreme much temperatures more extreme and temperatures have to be consideredand have to separately be considered later in separately the developm laterent in theprocess. development The internal process. battery The has internal more strictbattery thermal has more requirements strict thermal depending requirements on its current depending operational on its current state: Charging operational and state: dis- chargingCharging the and Li-Ion discharging batteries the are Li-Ion constraint batteries to a are tighter constraint margin to in a tightertheir operational margin in theirtem- peratureoperational range temperature in order to range prevent in orderdamage to preventor degradation. damage For or degradation.the used Panasonic For the 18,650 used ◦ cellsPanasonic charging 18,650 requires cells 0 charging to +45 °C requireswhile discharging 0 to +45 allowsC while −20 discharging to +60 °C, considering allows −20 the to ◦ +60anticipatedC, considering charge and the discharge anticipated currents charge provided and discharge by the currents power providedanalysis [17]. by the power analysisTypically, [17]. CubeSat architectures already provide a reasonable passive thermal design baselineTypically, which CubeSat can architecturesbe optimized already using suitable provide thermal a reasonable coatings passive as finishes thermal [18,19]. design Sincebaseline more which than can85% be of optimized the external using surfaces suitable are covered thermal by coatings solar cell as ﬁnishespanels, these [18,19 finishes]. Since aremore only than applicable 85% of the on externalthe end surfacesfaces and are to coveredsome degree by solar on cellthe remaining panels, these structural ﬁnishes ele- are ments.only applicable Because onof the the very end faceslow power and to dissipat some degreeion of on the the selected remaining integrated structural circuits elements. and otherBecause electrical of the components, very low power internal dissipation contributions of the selectedto the thermal integrated energy circuits balance and are other an- ticipatedelectrical to components, be relatively internal low. Notable contributions contributions to the thermal are only energy expected balance to are be anticipatedthe ADCS actuatorto be relatively systems low. (magnetorquers Notable contributions and reaction are wheels) only expected as well toas bethe the UHF ADCS transmitters. actuator Hence,systems placing (magnetorquers the battery andin the reaction center of wheels) the satellite as well near as the the ADCS UHF transmitters.and Tx/Rx hardware Hence, isplacing considered the battery advantageous. in the center However, of the moving satellite the near Up/Downlink the ADCS andhardware Tx/Rx away hardware from its is considered advantageous. However, moving the Up/Downlink hardware away from its current location at the end of the satellite near the antenna is currently unsuitable and has current location at the end of the satellite near the antenna is currently unsuitable and has to undergo additional considerations. Therefore, using an active heating system in order to undergo additional considerations. Therefore, using an active heating system in order to to keep the battery in its thermal operation range is being considered. The selected battery keep the battery in its thermal operation range is being considered. The selected battery pack already has an integrated autonomous heater system which will be used to heat the

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pack already has an integrated autonomous heater system which will be used to heat the battery if necessary [11]. Thermal analyses with ﬁnite element methods (FEM) and the thermal analysis software ESATAN-TMS (Thermal Modeling Suite) using the ﬁnal design of the satellite has to be conducted in order to verify the given thermal condition later in the development.

3.3. Power The power system is composed of three main components, i.e., a battery pack, solar panels and a PCDU. The purchased battery is designed for the use in nano-satellites and is composed of high-capacity lithium ion cells in a 2S-4P configuration with a total capacity of 10.4 Ah. It is equipped with a heater system, expanding the operational temperature range. The 3U solar panels, which are radiation and temperature resilient, consist of Azur Space 3G30A triple-junction solar cells with up to 30% efficiency, comprising a temperature and sun sensor. The PCDU is a contribution by the Institute of Space Systems of DLR Bremen and follows a new approach of high modularity, giving the ability to be tailored for different mission requirements by use of design automation techniques. It is set up to be fault–tolerant, due to redundancy and includes standard features as maximum power point tracking (MPPT) for battery charging or latch up protection of the outputs. It delivers 19 switchable output channels, one for each module on the satellite. The power system fulfils the power budget requirements see Section 2.3.

3.4. Attitude Determination and Control System—ADCS Given a 600 km sun-synchronous LEO, InnoCube is expected to be subject to external disturbance forces lower than 10−7 Nm, not considering the satellites residual magnetic moment. It is hard to estimate the residual magnetic moment, but literature analysis expects it to be less than 0.1 Am2 for a typical CubeSat [20,21]. Considering 0.2 Am2 torquer rods allows for an active magnetic actuation with a maximum torque of 1 × 10−5 Nm depending on the orbital position. System requirements imposed on the ADCS derive mainly from the EPISODE payload. The communication subsystem does not require specific attitude control because of the antenna’s omnidirectional characteristics. The EPISODE payload requires two main scientific attitude control modes: GNSS Antenna Pointing and Laser Ranging Ground Station Tracking. GNSS Antenna Pointing resembles an anti-nadir tracking, which ensures an unobstructed field-of-view into medium Earth orbit (MEO) of the GNSS antenna. Assuming a ±50 deg antenna beam width leads to a ± 130 deg anti-nadir attitude requirement. Assuming a 600 km circular orbit leads to maximum attitude of 3.75 deg/min for anti-nadir tracking. Laser ranging ground station tracking orientates the LRR in the direction of the laser ranging ground station and tracks this specific attitude. Figure 10 associates different contact elevation angles with required torque for optimum tracking performance. The LRR tolerates ±17 deg deviation from the optimum tracking orientation which is 20 deg off-axis pointing of optimum line-of-sight. With final simulation still pending, 2.3 mNm has been chosen to be the minimum torque requirement for the attitude actuation. In order to accommodate these requirements, a typical 3-axis ADCS will be implemented [18,19]: AerospaceAerospace2021 2021, 8,, 1278, x FOR PEER REVIEW 14 of 20 14 of 20

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Figure 10. 10. ExampleExample required required torque torque for foroptimum optimum pointi pointingng of different of different contact contact elevation elevation angles. angles. Figure 10. Example required torque for optimum pointing of different contact elevation angles. FigureFigure 11 11 gives gives a abasic basic system system overview: overview: attitude attitude determination determination will be will carried be carriedout out usingFigure redundant redundant 11 gives magnetic magnetic a basic and system andgyroscope gyroscopeoverview: sensors a sensorsttitude combined determination combined with six custom with will sixbe two carried custom axis sun- out two axis sensors, developed at the University of Wuerzburg. An extended Kalman filter will be sun-sensors,using redundant developed magnetic at and the gyroscope University sens ofors Wuerzburg. combined with An extendedsix custom Kalman two axis ﬁltersun- will be usedsensors, for sensor-datadeveloped atfusion the University and final attitude of Wuer eszburg.timation. An Thisextended allows Kalman an absolute filter attitude will be used for sensor-data fusion and ﬁnal attitude estimation. This allows an absolute attitude referenceused for sensor-data necessary for fusion the attitude and final controlling attitude es astimation. well as Thisthe estimation allows an ofabsolute the gyroscope attitude reference necessary for the attitude controlling as well as the estimation of the gyroscope biases.reference A yet necessary to be determined for the attitude quaternion controlling feedback as well controller as the estimationwill serve forof thethe gyroscopenadir and biases.trackingbiases. A A control yet yet to to be laws. be determined determined A basic B-dot quaternion quaternion bang bang feedback feedback detumbling controller controller controller will serve will can for serve be the used fornadir for the andthe nadir and trackinginitialtracking damping control control of laws. laws. the rotationalA Abasic basic B-dot rates B-dot bang as bang well bang as bang detumblingfor possible detumbling controllerdesaturation controller can of be the canused wheels. be for used the for the initialinitial damping of of the the rotational rotational rates rates as well as well as for as possible for possible desaturation desaturation of the wheels. of the wheels.

Figure 11. Block diagram of attitude determination and control system (ADCS) subsystem. Figure 11. Block diagram of attitude determination and control system (ADCS) subsystem. Figure 11. The lowBlock dynamic diagram GNSS of attitude antenna determination pointing mode and can control be accommodated system (ADCS) by subsystem. the use of threeThe magnet-torquer low dynamic actuators GNSS antenna only. Thepointing magn modeet-torquers can be will accommodated be based on bya self-devel- the use of The low dynamic GNSS antenna pointing mode can be accommodated by the use of opedthree ferritemagnet-torquer core torque actuators rod. Because only. The of themagn aforementionedet-torquers will minimum be based torqueon a self-devel- require- three magnet-torquer actuators only. The magnet-torquers will be based on a self-developed ments,oped ferrite laser coreranging torque ground rod. stationBecause tracking of the aforementioned requires the use minimum of a three-axis torque reactionrequire- ferritewheelments, coreactuationlaser torque ranging system. rod. ground BecauseThe magnet-torquersstation of the tracking aforementioned requiresare utilized the minimum touse desaturate of a torquethree-axis the requirements, wheels reaction if laser rangingwheel actuation ground system. station trackingThe magnet-torquers requires the are use utilized of a three-axis to desaturate reaction the wheelwheels actuationif system. The magnet-torquers are utilized to desaturate the wheels if necessary. The ADCS software includes the telecommand and telemetry management as well as fault detection and isolation (FDIR) procedures. Aerospace 2021, 8, x FOR PEER REVIEW 15 of 20

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necessary. The ADCS software includes the telecommand and telemetry management as well as fault detection and isolation (FDIR) procedures. TheThe ADCSADCS hardware (except (except for for the the external external sun-sensors) sun-sensors) will will be becombined combined in a inded- a dedicatedicated mechanical mechanical structure, structure, which which will will subsequently subsequently be mounted be mounted inside inside the satellite. the satellite. This Thisstructure structure also also includes includes the theADCS ADCS mainboard mainboard along along with with the themicrocontroller microcontroller running running the theADCS ADCS software. software. The advantage The advantage of this ofclosed- this closed-boxbox design is design the improved is the improved stand-alone stand- test- aloneing capabilities testing capabilities without the without need thefor a need complete for a completelyly assembled assembled satellite structure. satellite structure. Figure 12 Figureshows 12 theshows CAD-model the CAD-model (computer-aided (computer-aided design model) design model)of the ADCS of the ADCSbox which box whichcan be canintegrated be integrated into the into satellite the satellite using usingthe mounti the mountingng wedgelocks wedgelocks on the on upper the upper and lower and lower sides. sides.The only The electrical only electrical interface interface is the isback theplane-connector backplane-connector mounted mounted on the on mainboard. the mainboard.

Figure 12. Swappable ADCS box accommodating control electronics and actuators. Figure 12. Swappable ADCS box accommodating control electronics and actuators.

3.5.3.5. Command Command and and Data Data Management Management System System TheThe datadata managementmanagement systemsystem andand allall avionicsavionics consistconsist ofof a a network network of of redundant redundant SKITHSKITH modules. modules. The The OBCOBC isis thethe centralcentral element,element, responsibleresponsible forfor commanding,commanding, telemetry,telemetry, surveillance,surveillance, housekeeping, housekeeping, payload payload interfaces interfaces and and power power control. control. There There are four are instances four in- ofstances the OBC of the on theOBC satellite. on the Thesatellite. software The software will be based will be on based Rodos, on as Rodos, it offers as the it offers optimal the supportoptimal for support distributed for distributed and redundant and redundant computing. computing.

3.5.1.3.5.1. Onboard Onboard Computer Computer (OBC) (OBC) Hardware Hardware TheThe OBC OBC hardware hardware is basedis based on on the the SKITH SKITH interface, interface, using using the EFR32FG12 the EFR32FG12 for both fordata both processingdata processing and communication and communication with other with modules.other modules. Since theSince EFR32FG12 the EFR32FG12 microcontroller microcon- istroller powerful is powerful enough enough for most for applications, most applications, it will it not will only not deliveronly deliver the communication the communica- interface,tion interface, but will but also will serve also serve as a controller, as a controller, handling handling sensors, sensors, actuators actuators and memories. and memo- It impliesries. It aimplies 40 MHz a 40 Cortex-M4 MHz Cortex-M4 CPU with CPU 1 MB with ﬂash 1 MB memory flash andmemory 256 kB and RAM. 256 AkB property RAM. A thatproperty makes that the makes EFR32FG12 the EFR32FG12 perfectly suitableperfectly for suitable a small for satellite a small applicationsatellite application is its low is powerits low consumption power consumption with a maximum with a maximum power of power just 36.3 of mW just while36.3 mW sending while and sending 27.7 mW and when receiving data over the 2.4 GHz interface. Additionally, the OBC is equipped with an 27.7 mW when receiving data over the 2.4 GHz interface. Additionally, the OBC is external SPI 128MByte Flash and 1MByte MRAM, which have already been subjected to equipped with an external SPI 128MByte Flash and 1MByte MRAM, which have already total-dose radiation tests. been subjected to total-dose radiation tests. Figure 13 depicts a PCB, carrying two OBC modules. The layout is identical for both Figure 13 depicts a PCB, carrying two OBC modules. The layout is identical for both modules, but mirrored due to thermal reasons. Since the PCB dimension is given by the modules, but mirrored due to thermal reasons. Since the PCB dimension is given by the mechanical structure, this layout offers an area for additional components. mechanical structure, this layout offers an area for additional components.

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FigureFigure 13. 13.Overview Overview of of an an onboard onboard computer computer (OBC) (OBC) module. module.

3.5.2.3.5.2. Software Software TheThe SKITH SKITH hardware hardware is is designed designed in in a a modular modular way, way, having having different different modules modules for for differentdifferent tasks. tasks. For For the the software, software, the the same same modularity modularity principle principle is is used. used. Each Each SKITH SKITH mod- mod- uleule runs runs software software with with common common tasks tasks (e.g., (e.g., housekeeping) housekeeping) and and module-specific module-specific tasks. tasks. To To addressaddress modularity modularity in in software, software, we we follow follow an an app-centric app-centric approach approach for thefor softwarethe software design. design. 3.5.3. Onboard Software 3.5.3.The Onboard onboard Software software is based on two frameworks: Rodos and Corfu. Rodos is an open sourceThe onboard operating software system is withbased flight on two heritage frameworks: mainly developed Rodos and at Corfu. the University Rodos is ofan Wuerzburgopen source and operating the German system Aerospace with flight Center. heritage It ismainly able to developed run on different at the University processors of andWuerzburg platforms and in the bare-metal German execution,Aerospace butCenter. it also It is runs able as to a run guest on ondifferent other operatingprocessors systems,and platforms such as in Linux. bare-metal For bare execution, metal execution, but it also Rodos runs comes as a guest with aon real-time other operating preemptive sys- scheduling to adhere to timing requirements. A hardware abstraction layer allows to tems, such as Linux. For bare metal execution, Rodos comes with a real-time preemptive develop on a desktop computer and directly deploy the code to the target platform. scheduling to adhere to timing requirements. A hardware abstraction layer allows to de- Besides the low-level part, Rodos comes with a sophisticated communication middle- velop on a desktop computer and directly deploy the code to the target platform. ware. The middleware applies the publish/subscribe principle. Users instantiate topics, Besides the low-level part, Rodos comes with a sophisticated communication mid- which publishers use to publish new data, and subscribers register for notification on dleware. The middleware applies the publish/subscribe principle. Users instantiate topics, data delivery. Beyond the local communication, the middleware comes with gateways which publishers use to publish new data, and subscribers register for notification on data that enable publishing and subscribing topics across several computing nodes (i.e., Rodos delivery. Beyond the local communication, the middleware comes with gateways that en- instances). Rodos middleware is used for exchanging messages via radio between the able publishing and subscribing topics across several computing nodes (i.e., Rodos in- SKITH modules. stances). Rodos middleware is used for exchanging messages via radio between the While Rodos provides the basic infrastructure for executing software on hardware SKITH modules. and providing a communication middleware, the Corfu framework provides a common structureWhile for Rodos onboard provides software. the Developersbasic infrastructu definere topics, for executing apps, and software nodes on formally hardware in configurationand providing (specification) a communication files. A middleware, node represents the theCorfu instance framework of a software provides that a common runs on astructure SKITH module for onboard a selection software. of apps. Developers define topics, apps, and nodes formally in configurationApps are a(specification) collection of files. threads A node and represents subscribers the that instance accomplish of a software a specific that task, runs foron examplea SKITH housekeeping,module a selection attitude of apps. control, or time management. Each app defines a communicationApps are a interface collection towards of threads other and applications subscribers and that towards accomplish the grounda specific segment. task, for Developersexample housekeeping, describe the communicationattitude control, interfaceor time management. between the appsEach byapp defining defines whicha com- topicsmunication the apps interface publish towards to and which other applicatio they subscribens and to. towards The wiring the ofground the topic segment. publication Devel- andopers subscription describe the is accomplishedcommunication by interface the node. between the apps by defining which topics the Theapps app publish interface to and towards which thethey ground subscribe segment to. The comprises wiring of a the list topic of telecommands publication and to whichsubscription the app is reacts accomplished and a list by of the telemetry node. that the app transmits. A special case is the standardThe (real-time)app interface telemetry. towards While the ground apps transmit segment normal comprises (extended) a list of telemetry telecommands only on to request,which the the app standard reacts telemetry and a list is of transmitted telemetry periodically.that the app Thetransmits. standard A telemetryspecial case gives is the a standard (real-time) telemetry. While apps transmit normal (extended) telemetry only on

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quick overview of the node’s vital status. Therefore, every app should contribute some variables with the most important vital parameters. Based on the information of the configuration files, Corfu generates code with abstract methods that the developers have to implement with their intended behaviour. Corfu already covers the communication handling—either in its library or in the generated code. Methods just have to be implemented that pass the parameters in a plain structure or invoke methods to send data. To provide rapid prototyping of onboard software, Corfu also comes with a generic ground software. It uses the same information of the configuration files to show a graphical user interface for sending telecommands and printing telemetry data.

3.5.4. InnoCube Apps As outlined in the former subsection, all apps are developed in Corfu and each app can be configured to fit different nodes, before being compiled into a specific Rodos image. A set of vital apps are present on all nodes enabling basic overview and operations, e.g., the AnomalyReporter, Housekeeper, Surveillancer, and Watchdog, which are continuously monitoring and reporting on the state of each node. Furthermore, there are the apps needed to upload operating system images, write them to flash memory and boot the node, namely the FirmwareUploader, FlashManager and BootManager. Finally, there is always a TimeManager client present for CubeSat-wide time synchronisation. There are six different configurations of SKITH nodes that come with the following apps in addition to the vital apps mentioned above: The OBC node software features the Commander, which receives the telecommands sent from ground segment, verifies their signatures and publishes them to the Telecommand topic. The more complex TimedCommander stores and distributes lists of time-tagged commands, which can also be edited via telecommand before activation. Reacting on specific telecommand lists for certain satellite modes, the Navigation and Pow- erManagement apps issue high level commands to the SKITH nodes which are controlling the respective subsystems. The PCDU node image includes the BatteryVoltageManager, the PowerManagementExecutor implementing the commands issued by the PowerManage- ment as well as the RedundancyManager ensuring that only one of each hot redundant SKITH nodes is taking lead. In addition, this image comprises a MasterWatchdog and the SKITHRadioMaster, which issues communication frames to each node. The up/downlink node is given the Uplink and Downlink apps needed to interact with the transceiver, as well as the DownlinkManager, which manages different queues of telemetry to be sent down and the TelemetryHistory app, holding older telemetry to be sent down when the link is otherwise unused. The attitude control node comes with the AttitudeDetermination and AttitudeControl apps, including the FDIR apps, which are explained in the ADCS section. The payload computers for EPISODE and Wall#E will have dedicated apps to interact with each payload and forward the respective telecommands and telemetry.

3.5.5. Reporting and Command Flow All SKITH nodes within InnoCube communicate over radio using the Rodos middleware on top of the radio protocol. The middleware topics and publish/subscribe mechanism provide a structured means of onboard communication. Figure 14 depicts the ﬂow of information between the apps on InnoCube’s SKITH nodes as facilitated by the Rodos middleware: telecommands, depicted in orange, are sent from ground and received by the Uplink app, which publishes them to a topic only the Commander app on the OBC subscribes to. Aerospace 2021, 8, 127 18 of 20 Aerospace 2021, 8, x FOR PEER REVIEW 18 of 20

FigureFigure 14. 14.Telecommand, Telecommand, telemetry telemetry and and anomaly anomaly ﬂow flow between between apps apps via via the the Rodos Rodos middleware. middleware.

AfterAfter verifyingverifying thethe signature of of the the telecommand, telecommand, the the CommanderCommander publishespublishes any any tel- telecommandecommand to to be be executed executed immediately immediately to the Telecommand topic.topic. InIn case case the theCommander Commander receivesreceives a a timed timed telecommand telecommand to beto executedbe executed later, la itter, publishes it publishes it to theit to TimedTelecommand the TimedTelecom- topic,mand from topic, where from the whereTimedCommander the TimedCommanderapp retrieves app andretrieves stores and it. When stores the it. telecommandWhen the tel- isecommand due for execution, is due for the execution,TimedCommander the TimedCommanderpublishes it publishes to the Telecommand it to the Telecommand topic. All appstopic. subscribe All apps to subscribe the Telecommand to the Telecomma topic andnd try topic to execute and try every to execute command every received command via thisreceived middleware via this topic middleware that was topic addressed that was to themaddressed and isto formed them and correctly. is formed Malformed correctly. orMalformed unknown or telecommands unknown telecommands are dropped are by dro thepped app by and the a non-criticalapp and a non-critical anomaly (black anom- arrow)aly (black is published arrow) is topublished the anomaly to the topic. anomaly If a severetopic. If error a severe occurs, error apps occurs, can alsoapps publish can also criticalpublish anomalies. critical anomalies. The ModeManager The ModeManageron the OBC on listens the OBC for anomalies listens for and anomalies switches and to safeswitches mode to depending safe mode on depending their severity. on their In save severity. mode In the saveTimedCommander mode the TimedCommanderpublishes a setpublishes of telecommands a set of telecommands from its save from mode its save list to mode the Telecommand list to the Telecommand topic. There topic. are There two kindsare two of telemetry,kinds of telemetry, both depicted both depicted in green: in the green: standard the standard telemetry telemetry of a node of a comprises node com- ofprises a set of of a telemetryset of telemetry of all of of all its of apps its apps that that published published periodically periodically by by the theHousekeeper Housekeeper appapp to to thethe StandardTelemetryStandardTelemetry topic.topic. Extended telemetry telemetry can can be be enabled enabled and and published published by bytelecommand telecommand for for each each app. app. Both Both the theDownlinkManagerDownlinkManager andand the theTelemetryHistoryTelemetryHistory appapp sub- subscribescribe to the to the telemetry telemetry channels. channels. The The DownlinkManagerDownlinkManager schedulesschedules the thetelemetry telemetry to be to sent be sentdown down and andforwards forwards it via ita vianode a local node topic local to topic the Downlink to the Downlink app, whichapp, uses which the usestransmit- the transmitterter to send toit to send ground. it to ground.The TelemetryHistory The TelemetryHistory stores olderstores telemetry, older telemetry, which the which Downlink- the DownlinkManagerManager will try towill schedule, try to schedule, at a lower at apriority, lower priority, when the when link theis unused. link is unused. 3.6. Communication 3.6. Communication The main focus of the communication system is reliability and simplicity, as the The main focus of the communication system is reliability and simplicity, as the tech- technological demonstration of the satellite shall not be endangered by a failure of the nological demonstration of the satellite shall not be endangered by a failure of the com- communication link. The scientific data generated by the satellite consists of telemetry data munication link. The scientific data generated by the satellite consists of telemetry data and positional data, therefore high data rates are not necessary. and positional data, therefore high data rates are not necessary. As first step of the design process, the frequencies for uplink and downlink are defined. As first step of the design process, the frequencies for uplink and downlink are de- The mission consists of education and technology demonstration. Consequently, amateur fined. The mission consists of education and technology demonstration. Consequently, frequencies, which offer a number of advantages of a financial and organizational nature, areamateur being used.frequencies, which offer a number of advantages of a financial and organizational nature,For simplicityare being used. and budget reasons, only one antenna can be included. As a result, only one bandFor willsimplicity be used and as well:budget UHF reasons, frequencies only on ine the antenna range ofcan 435–438 be included. MHz. As a result, onlyThe one communication band will be used is performed as well: UHF by commerciallyfrequencies in available the range transceivers. of 435–438 MHz. A tradeoff lead toThe the communication selection of the is RadioBro,performed MiniSatCom by commercially UHF availabl transceiver.e transceivers. Although A lackingtradeoff extensivelead to the flight selection heritage, of the its size,RadioBro, power MiniSa consumption,tCom UHF and transceiver. configurability Although and comparably lacking extensive flight heritage, its size, power consumption, and configurability and comparably

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low cost lead to its use in the mission. With this product, two identical flight units can be integrated into the satellite. A redundant configuration is implemented, with two identical transceiver units on separate boards. As both radios share a single antenna, a radiofrequency (RF) switch is used to change between the units. Multiple layers of internal identification are integrated to determine whether a transceiver is operable or not. Each of the transceiver boards additionally consist of two SKITH nodes, eliminating any single-point-of-failure. The satellite design utilises only a single frequency antenna to avoid two antennas interfering with each other. As a result, the channel is capable of half duplex communication. Different configurations of antennas have been analysed, with regard to commercially available products and the UHF antenna module of EnduroSat has been selected. It represents a trade-off between cost, radiation pattern and deployment mechanism. The maximum gain is about 0 dBi, although the omnidirectional pattern is more important than its peak gain.

4. Conclusions Outlook Aboard the InnoCube mission of TU Braunschweig and the University of Wuerzburg, the next level of innovative technologies will be flown. The first wireless satellite data bus using SKITH technology should provide numerous advantages that are utilized in the 3U CubeSat design. An easily accessible modular structure enables the switching of satellite modules within minimum time and, theoretically, in any phase of the project. Hardware and software are structured in a modular way, having different modules for different tasks and following an app-centred approach based on two frameworks: Rodos and Corfu. In addition, a novel experimental structural energy storage WallE-2-Space payload is to be tested and the performance analysed in a relevant operational environment. The development of a software-defined GNSS receiver that allows a flexible selection of GNSS signals should enable precise orbit knowledge for small satellites. This makes conjunction assessments more accurate, can prevent collisions, may help to stabilize the space debris environment and stimulates future research in this field. To verify the GNSS accuracy, a laser retro reflector is designed and scaled down to fit to a CubeSat system. The upcoming steps are the manufacturing, integration, and test in the given timeline until the envisioned launch in early 2023 and verification of the technologies within the minimum one year in orbit operation. The upcoming challenges of this project are the detailed electronic board design and experimental charge regulation of the WallE-2-Space payload and finding the best integration scenario that minimizes the effect on battery performance. Furthermore, the SKITH modules need to be integrated into the circuit board design of the payload PCB and modules with COTS components. Finally, the software must be completed, especially the payload software and payload configuration app. All those tasks within this CubeSat project will create an attractive, practice oriented and practice relevant form of education for the involved students who will acquire specific technical knowledge and learn at an early stage to work in a team and result-oriented manner according to deadlines.

Author Contributions: Conceptualization, S.M., E.S. and B.G.; software F.S. and F.F.; formal analysis, J.-L.K. and M.T.; investigation E.D. and S.G.; writing—original draft preparation, B.G.; writing— review and editing, S.G., B.G., T.B., T.W., F.F., F.S., E.D. and E.S.; visualization, M.T. and J.-L.K.; project administration, T.W. and B.G.; funding acquisition, S.M. and E.S. All authors have read and agreed to the published version of the manuscript. Funding: This work is supported (support code 50RU2001/50RU2000) by the Federal Ministry for Economic Affairs and Energy on the basis of a decision by the German Bundestag. We acknowledge support by the German Research Foundation and the Open Access Publication Funds of Technische Universität Braunschweig. Acknowledgments: The authors like to acknowledge the great cooperation and work of the Avionics Systems department of Institute of Space Systems of DLR Bremen. Thanks to the Institute for Particle Aerospace 2021, 8, 127 20 of 20

Technology (iPAT) of Technische Universität Braunschweig for previous and continuous cooperation and support with Wall#E. Further, the authors thank Ludwig Grunwaldt who provided the testbed and his expert knowledge for testing the prisms for the laser ranging payload at the GFZ Potsdam. Also, thanks to our sponsors Valispace and ITP Engines UK Ltd. for ESATAN-TMS. We further acknowledge the motivated students, who started to form the CubeSat group and are going to work and accompany us during this project until the launch and beyond. Conﬂicts of Interest: The authors declare no conﬂict of interest.

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