ESTCUBE-2 MISSION AND SATELLITE DESIGN Hendrik Ehrpais(1,2), Indrek Sünter(1,2), Erik Ilbis(1,2), Janis Dalbins(1,2), Iaroslav Iakubivskyi(2), Erik Kulu(2,4), Indrek Ploom(1,2), Pekka Janhunen(3), Joel Kuusk(1), Jānis Šate(5), Roberts Trops(5), Andris Slavinskis(1,2,3) (1) Tartu Observatory, Observatooriumi 1, 61602 Tõravere, Estonia, +372 696 2510, [email protected], [email protected], [email protected], [email protected], [email protected], [email protected], [email protected] (2) University of Tartu, Institute of Physics, Tähe 4-111, 51010 Tartu, Estonia +372 737 6524, [email protected] (3) Finnish Meterological Institute, Erik Palmenin aukio 1, P.O. Box 503, FI-00101, Helsinki, Finland, [email protected] (4) Radius Space, Akadeemia 21/1, Tallinn, 12618, Estonia, [email protected] (5) Ventspils University College, Inzenieru 101, LV-3601, Ventspils, Latvia, [email protected] , [email protected]

ABSTRACT Here we present the preliminary mission design for the ESTCube-2 three-unit CubeSat. Its main mission is to test Coulomb drag propulsion. Coulomb drag can be used in Low-Earth Orbit by deploying and charging a tether that is used to brake the orbital velocity of the satellite and reduce its orbital altitude. To test this concept, ESTCube-2 will deploy and charge a 300 m tether. Such a tether could deorbit ESTCube-2 from the altitude of 700 km to 500 km in half a year. Other payloads that are being considered for the ESTCube-2 satellite are an Earth observation camera, a C-band communications system and an experimental laser communication system. ESTCube-2 in-orbit demonstration platform will also be designed for other electric solar wind sail experiments outside of the influence of Earth’s magnetic field. The satellite bus will be integrated into one system that could also be reused for different types of missions. The integrated system is developed to maximise the space for payloads on a nanosatellite. This paper presents the payloads and system design of ESTCube-2. 1

INTRODUCTION

ESTCube-2 is a nanosatellite developed mostly by the students of University of Tartu and students that join the ESTCube program from all over the world. The satellite builds on the successful mission of ESTCube-1 [1] with the aim to develop a small and competent satellite bus solution. ESTCube-2 is a 3-unit CubeSat built according to the CubeSat standard[2]. The satellite aims to further test Coulomb drag propulsion [3] that was planned to be tested on ESTCube-1 and is going to be tested this year on Aalto-1 [4]. The satellite is also designed to be able to test an Earth observation payload and a laser communication payload. These payloads are supported by high speed communications [5] and a cold gas thruster payload [6]. The satellite is designed with the approach to integrate as many of the satellite subsystems as possible to reduce the size of the satellite bus. The development of the satellite is divided into the following subsystems: electrical power system (EPS), communication subsystem (COM), on-board computer (OBC), attitude and orbit control system (AOCS) and structure (STR). The satellite is expected to be completed in 2018.

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MISSION ANALYSIS

The current focus of the team is building a satellite, that could accommodate a variety of payloads, which are being considered on the ESTCube-2 satellite. Those payloads are described below. 2.1

Coulomb Drag Propulsion

The plasma brake is a lightweight, efficient, cost-effective and scalable deorbiting system with a potential to address the space debris problems in the most critical altitudes of 900 km and less [7,8]. The experience with ESTCube-1 and Aalto-1 has shown that it is feasible to host a similar payload in ⅓ to ½ of a CubeSat unit. ESTCube-2 will deploy and charge a 300 m tether, which will be used to reduce the orbit altitude of the satellite. The negatively charged plasma brake tether interacts with the ambient ionospheric plasma ram flow to slow down the satellite. Such a tether could deorbit ESTCube-2 from the altitude of 700 km to 500 km in half a year. The mass of such a long tether is 30 grams according to a conservative estimate. The main requirements for the satellite bus are to provide the total angular momentum of 23 Nms for centrifugal tether deployment, to provide means of deployment verification and to provide up to 3000 mW of power for the payload. A NanoSpace MEMS-based cold gas micropropulsion system will be used and tested on ESTCube2 for a potential follow-up mission that would test the electric solar wind sail(E-sail) outside of Earth’s magnetic field [6]. The payload is developed by the Finnish Meteorological Institute. 2.2

Earth Observation Payload

The Earth observation payload on ESTCube-2 could be a technical demonstration mission for miniaturised remote sensors on small satellite for various Earth observation missions. The aim of the payload is to take multispectral images in specific wavelengths for scientific purposes. The payload would use one common lens, two sensors and optical beam splitting to take multispectral images. The payload would take up to 1U of space on ESTCube-2 and weigh less than 1 kg. This payload sets considerable requirements for the data transmission and storage for the satellite. A sufficient amount of data transfer would be up to 200 MB per day and storage of up to 0.5 GB of images. The most important requirement for the attitude control of the satellite is to minimise the rotation as much as possible when taking images and track the targets. The payload is being investigated by Tartu Observatory. 2.3

High Speed Communications Payload

One of the ESTCube-2 payloads is going to be an experimental high speed communication subsystem developed by Ventspils University College. The system is based on a FPGA architecture, which will provide programmable changes in data-rate and modulations during downlink period. The system is going to use phase modulations, like BPSK, QPSK, 4APSK, 8APSK, 16APSK, 32APSK. The lowest data-rate for this system is planned to be 1 Mbps. Rectangular microstrip (Patch type) antenna will be used to transmit data in radio amateur C-band frequency range. Exact downlink frequency will be determined after frequency coordination process through the International Amateur Radio Union and the International Telecommunication Union. In case of successful tests of this payload system will be used to transmit data from satellite with higher datarate than possible in the main COM [5]. 2.4

Laser Communications Payload

Under consideration is an optical space-to-ground communications system called LaserCOM for short. Very preliminary requirements are 0.3U volume and external laser emitter aperture not more The 4S Symposium 2016 – H. Ehrpais

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than 10 mm. Power consumption during communication sessions could be even up to 50 W for few minute long periods like for AeroCube-OCSD [9], which probably requires separate high-current batteries or supercapacitors. Waste heat dissipation will be one of the challenges for the structure and thermal control when that high power consumption is needed. For this mission the pointing control accuracy should be aimed to be within 0.1 degrees and on-board GPS is needed for precision orbit determination to achieve positional errors less than 10 meters. This payload is investigated by the University of Tartu. 3

SATELLITE DESIGN

The satellite bus will be built completely in-house by students and experienced individuals from the ESTCube-1 satellite mission. All of the subsystems present in the satellite bus are developed so that the volume of the integrated system could be minimised. The satellite architecture is described in Figure 1. The electrical power system and the communications system will jointly use the same MSP430FR series microcontroller unit (MCU). The rest of the subsystems are run by the main MCU, STM32F7. The system also has intelligent side panels, which hold a MSP430FR MCU, perform maximum power point tracking for solar panels and also hold sensors for the AOCS system. The integrated bus solution is illustrated in Figure 2. The current design of the integrated bus takes 0.5 U of space in the satellite. It is also possible to perform firmware updates on all of the MCUs after the launch of the satellite. This enables the team to correct any unexpected problems that the satellite may encounter. A similar firmware update system was also used in ESTCube-1[10].

Figure 1. Satellite architecture drawing The 4S Symposium 2016 – H. Ehrpais

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Figure 2. Satellite bus illustration 3.1

Structure

Taking into consideration lessons learned from the ESTCube-1 mission, a monoblock structure is not suitable for our future CubeSat missions. A multiple panel frame is used in the ESTCube-2 satellite. Such an approach increases the accessibility and simplicity of design and decreases the time and cost of manufacturing. Satellite structure will consist of a primary and secondary frame made from the aluminium alloy 7075-T6. The primary frame includes two mirrored panels fastened with auxiliary ribs. The main printed circuit boards (PCBs), reaction wheels, batteries and a star tracker have been housed inside a secondary frame. In addition, the aluminium PCB side panels will contain solar cells and sun sensor. The star tracker baffle, located almost along the axis of rotation, extends for approximately 8 mm along the surface normal. This is below the maximum allowed by ISIPOD CubeSat deployer [11]. Spacecraft mechanisms are mainly related to deployable solar panels and deployable low speed communication antennas. Nylon or polyester core will be cut using a heated up nichrome wire, which will ensure separation. In order to successfully deploy the antenna, it will use a suppressed copper strip. The deployable panels will be using hinges. In the initial position two deployable panels will extend 6 mm-s along the surface normal. The structure with deployed solar panels is illustrated in Figure 3.

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Figure 3. Satellite drawing with deployed panels 3.2

Attitude and Orbit Control System

The attitude and orbit control system is designed with the aim to provide enough angular momentum for the centrifugal deployment of the E-sail tether. The AOCS system is designed to spin up the satellite to one revolution per second while aligning the spin axis with Earth’s polar axis with a pointing error of less than 3 degrees. The attitude determination is designed to also fulfil the requirements of a laserCOM and Earth observation system, which means that the required determination and pointing accuracy is under 0.1 degrees. Attitude determination is handled by magnetometers, Sun sensors, accelerometers, gyroscopic sensors and a star tracker. Attitude is controlled by magnetic rods and reaction wheels. The AOCS system is designed to also be used outside of LEO. For this reason, the AOCS algorithms are able to function without the magnetometers and magnetic torquers and are able to also use thrusters for attitude control. The star tracker and Sun sensors are developed in-house and the magnetometers, gyroscopes and the accelerometer are based on commercial off-the-shelf components. The reaction wheels from Hyperion Technologies are under consideration. The calculations of the AOCS system are performed by the main MCU, the star tracker will have a separate MCU for its calculations. 3.3

On-Board Computer

Both the electronics and software of the AOCS will be tightly integrated with the on-board computer. In addition to running AOCS algorithms, the main MCU (STM32F7 series) shall control the on-board payload modules, handle telemetry and telecommands. The OBC will enable timebased scheduling of on-board operations. For ESTCube-1, in-orbit firmware updates and on-board Pawn scripts enabled experimentation with various on-board algorithms and applications [10,12]. By improving the modularity of firmware images and introducing support for dynamic partial The 4S Symposium 2016 – H. Ehrpais

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firmware updates, the team plans to replace on-board scripting languages. This would allow for the management of complex in-orbit procedures with the performance of native code. The firmware modules would be stored in non-volatile random-access memory together with configuration tables. The OBC shall collect telemetry from the always-on-system, sidepanels, star tracker, AOCS sensors and from itself. After compression, the data will be stored in the on-board mass storage. In order to improve the compression ratio, it is planned to apply compression algorithms dedicated to specific types of data such as temperature, current or voltage. 3.4

Electrical Power System

The electrical power system is responsible for energy harvesting, energy storage and power distribution. The satellite will have solar cells on all four 3U sides and in addition, two deployable panels. The energy harvesting is handled by a maximum power point tracking system connected directly to the main power bus. The energy will be stored in a Li-Ion battery pack with two prismatic cells in series. The amount of battery packs is not limited by the electronics and will eventually depend on the payload requirements. Battery protection circuitry will protect the batteries automatically from over- and undervoltage conditions and from excessive charge and discharge currents. The voltage conversion system will provide regulated 3.3 V to the bus through current-limited switches. Where possible, battery bus will be used through current-limited switches in order to minimize losses and improve efficiency. The payloads will not be provided with the regulated buses due to the excessive conductor lengths within the satellite and if needed, will do their own regulation from the battery bus. 3.5

Communication Subsystem

The ESTCube-2 Communication subsystem is going to have a two-way half-duplex main communication system with uplink and downlink on 70 cm amateur radio band (frequency range 435-438 MHz). Main communication system is going to use integrated transceiver circuit Si4463 from Silicon Labs, which will provide changeable on-air baud rates from 9600 to 38400 bps with binary Gaussian frequency-shift keying (2GFSK). Both up- and downlink is going to support standard AX.25 unnumbered information frames as a transport protocol with AX.25 9600 baud radio amateur mode. Same Silicon Labs integrated circuit will also be used to provide on-off keying modulated signal output on a 70 cm amateur radio band for continuous wave beacon signal to transmit telemetry data in Morse code, which is robust and easy to receive and decode. Quarter wave monopole is planned to be used as an antenna for both up- and downlink. For downlink, an antenna impedance matching circuit will be used to achieve maximum transmission efficiency. A power amplifier on the satellite will provide up to 1 W for data downlink, and Si4463 transceiver circuit is capable of receiving a signal at up to -110 dBm, depending on data-rate and modulation type . An external low-speed Communication receiving system is also developed, which will ensure partial redundancy in case of firmware or hardware failure. This independent system will be based on integrated transceiver CC1310 with integrated MCU from Texas Instruments. This system will be capable of receiving GFSK modulated signal with data-rate up-to 50 kbps. This system will have a separate 70 cm band antenna for independent reception of commands. 4

CONCLUSIONS

In this paper, we presented the design of the ESTCube-2 nanosatellite. The main mission of the ESTCube-2 satellite is testing the Coulomb drag propulsion technology for de-orbiting the satellite. Other potential payloads include a Earth Observation camera and a LaserCOM communications system. The information from the payloads will be transported to the ground by a C-band high speed communications system. The 4S Symposium 2016 – H. Ehrpais

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The ESTCube-2 satellite uses a novel design approach by combining many of the subsystems in one small integrated solution to reduce the volume of the satellite bus. The article describes the design of the EPS, COM, AOCS, OBC and STR subsystems. The development of the subsystems and the integrated bus has already begun and the testing of the system is to be finished in 2018. 5

ACKNOWLEDGEMENTS

We are grateful to all ESTCube-2 team members and everyone involved in the project. The author would also like to thank Kristjan Jaak Scholarships. 6

REFERENCES

[1] Lätt S., Slavinskis A., Ilbis E., et al. ESTCube-1 nanosatellite for electric solar wind sail inorbit technology demonstration, Proc. Estonian Acad. Sci., 200-209, 2014 [2] CubeSat Design Specification Rev. 13, The CubeSat Program, Cal Poly SLO, California, 2014 [3] Janhunen P. and Sandroos A., Simulation study of solar wind push on a charged wire: basis of solar wind electric sail propulsion, Ann. Geophys., 25, 755–767, 2007 [4] Praks J., et al. Aalto-1 - An experimental nanosatellite for hyperspectral remote sensing, 2011 IEEE International Geoscience and Remote Sensing Symposium, IGARSS, 2011 [5] Sate J., Trops R., et al. Concept of the spectrally efficient CubeSat communication subsystem, Space review, Vol. 4, 2016 [6] Grönland T. A., et al. Development of MEMS-Based Spacecraft Propulsion, 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2006 [7] Janhunen P., Electrostatic Plasma Brake for Deorbiting a Satellite, Journal of Propulsion and Power, 26, 370-372, 2010 [8] Janhunen P., Simulation study of the plasma brake effect, Ann. Geophys., 32, 1207-1216, 2014 [9] Janson S. W., Welle R. P., The NASA Optical Communication and Sensor Demonstration Program, Proceedings of the 27th AIAA/USU Conference, Small Satellite Constellations, 2013 [10] Sünter I., Vahter A., et al. Firmware updating systems for nanosatellites, IEEE Aerospace and Electronic Systems Magazine, 2016 [11] ISIS, ISIPOD CubeSat Deployer Product Specification [12] Laizans K., Sünter I., et al. Design of the fault tolerant command and data handling subsystem for ESTCube-1, Proc. Estonian Acad. Sci., 222-231, 2014

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