OPTOS (Optical Nanosatellite)
OPTOS is a low-cost triple-cube nanosatellite project of INTA (Instituto Nacional de Tecnica Aerospacial), the Spanish Space Agency, Madrid. The overall objective is to demonstrate new technologies in spacecraft development such as: a distributed OBDH (On-Board Data Handling) subsystem based on FPGAs (Floating Point Gate Arrays), and CPLDs (Complex Programmable Logic Devices), an optical wireless communication system (OBCom) with a reduced CAN (Controller Area Network) protocol is implemented, the internal structure is based on composite materials. 1) 2) 3) 4) 5) 6) 7) 8)
Figure 1: Illustration of the OPTOS nanosatellite with solar panels sun sensors and antennas deployed (image credit: INTA)
The nanosatellite is designed and developed at INTA by using the ECSS (European Cooperation for Space Standards) of ESA including the qualification tests; every unit is certified and rigorously tested. The goal is to implement state-of-the-art technology into a small platform. The requirements call for a mission of 1 year duration using a platform that adheres to the CubeSat standard. Attitude control of the spacecraft is needed for the operation of an imaging camera.
The nanosatellite external structure corresponds to the triple CubeSat format (3U) in size and mass, i.e. 10 cm x 10 cm x 34.5 cm, and ~ 3 kg. Internally, a composite structure is selected to support all the elements and to allow easy integration and test. The spacecraft structure features an external aluminum casing provided by Pumpkin Inc. of San Francisco, CA, and an internal carbon fiber structure, design and manufactured at INTA. 9) 10) 11)
Figure 2: Axes references if the OPTOS spacecraft (image credit: INTA)
ADCS (Attitude Determination and Control Subsystem): The spacecraft is 3-axis stabilized. The attitude is sensed using 2 sun sensors (of TNO) which are mounted on the -Y and -Z faces of the spacecraft. The magnetic field is measured with a 3-axis fluxgate magnetometer developed at INTA. This device represents an improved design based on the HMC-1043 of Honeywell. In addition, a sun presence sensor, an alert device, is placed on the +z axis; it is used to warn the camera in case a sun incidence is occurring in this region.
Actuation is provided by a reaction wheel of AstroFein (Astro und Feinwerktechnik Adlershof GmbH,), Germany. It provides inertial pointing during observation and nominal modes of the spacecraft. Five magnetorquers of Clyde Space, UK/Scotland, (2 on x, 2 on y, and 1 on z-axis) provide freedom of rotation, with variant torque vector depending on time of actuation and sign of voltage used (±). The magnetorquers are used for attitude transitions (observation-navigation) and to desaturate the reaction wheel; they are being embedded in the solar panel’s PCBs (Printed Circuit Boards), which offers a considerable space saving inside the spacecraft. - The ADCS control algorithms have been developed at SENER, Spain.
Figure 3: Schematic view of the S/C body axis with respect to the sun vector (image credit: INTA)
EPS (Electrical Power Subsystem): Surface-mounted triple-junction GaAs are being used (1 string with 6 cells x 4 sides) as energy source, and a Li-ion battery for energy storage. - To receive an optimum amount of power during nominal (no observation) mode operations, the S/C will be oriented with its z-axis pointed perpendicular to the ecliptic and contained in orbit plane. This orientation will provide a power generation of 7.2 W at EOL. The power is distributed at the battery charge regulator board (Clyde Space Ltd., Glasgow, UK). The following regulated voltages are used by the subsystems: +3.3 V, +4 V, +5 V, +5.5 V, ± 12 V, and unregulated. The EPS uses DC/DC converters to distribute the regulated power to the different subsystems and payloads.
OBCom (On-Board Communications) subsystem: OBCom, of OWLS heritage, comprises an internal wireless optical communications subsystem, and a wireless CAN (Controller Area Network) bus is employed as the main TM/TC bus for the satellite; it is mandatory for all the units.
This implies that there is no need of cable usage between boards to communicate data, as almost every board is furnished with an emitting and a receiving diode that can pulse light at different frequencies into an optical channel, as well as reading the signals coming from other boards. These units are called OBCom modules and are independent intelligent units based on CPLDs (Complex Programmable Logic Devices). The configuration uses CAN drivers, derived from a CAN core implementation of ESA; hence, the on-board communications protocol is a reduced CAN bus customized by INTA. - The selected emitter is the SFH4205 which is duplicated in every module. The receivers are two TEMD5110 photodiodes and an IR filter to avoid visible-light interference. Photodiodes generate low level currents that are proportional to the level of illumination.
Each unit of the OBDH has an associated OBCom module. The module gives optical wireless communication capabilities to the OBDH through a CAN bus. The modules have been optimized in terms of budgets:
- Size: 25 mm x 15 mm x 14 mm
- Mass: 8 gram
- Power consumption: 50 mW.
Figure 4: Illustration of the OBCom module (image credit: INTA)
The transmission rate is 125 kbit/s. The optical on-board communication design is indeed a neat solution since it allows an easy integration of all subsystems (there is no need to assemble and disassemble cables).
OBDH (OnBoard Data Handling) subsystem: The distributed OBDH architecture is based on the CAN (Controller Area Network) protocol using CPLD/FPGA with a Microblaze processor to run the high level software. The architecture of OBDH is based on two types of interface units:
• EPH (Enhanced Processing Unit): This device provides service to a FPGA Xilinx VIRTEX-II ProTM that includes the TTC processor. The on-board software controller, the ADCS and TTC software run on the EPH. It has an interface with the TTC subsystem for receiving TC frames and subsystem status information, as well as for sending TM frames. It also interfaces with the rest of satellite through the CAN bus.
• DOT (Distributed OBDH Terminal): This interface device is based on a CPLD Cool Runner II implementation providing subsystem control of the platform (PDU, ADCS, thermal sensors) and payloads. A CAN bus interface is included for receiving commands or for sending data. By default they provide:
- Sixteen digital lines that can be configured as discrete outputs or clock signals
- Three analog channels and one AD convertor for 10 bit analog-digital conversion.
All the DOTs work by switching on and off periodically to save energy and to minimize SEUs (Single Event Upset); however, the devices remain switched on continuously when they detect some requests or when they need to transmit their data.
The on-board software controller, installed on the EPH, is in charge of managing the subsystems and payloads. It includes: CAN drivers, TTC drivers, the operating system kernel, ADCS software, and the application software.
Figure 5: Block diagram of the OBDH /OBCom subsystems (image credit: INTA)
TCS (Thermal Control Subsystem): Passive thermal control is implemented based in the selection of materials, internal arrangement and paints or MLI (Multi-Layered Insulation).
RF communications: Use of UHF (435 MHz) bands with transmissions in both directions. Use of a TNC (Terminal Node Controller) and 4 monopole antennas with omnidirectional radiation capability. The transponder works in half-duplex fashion using Manchester pulses (SP-L) for the downlink, and a phase with data subcarrier (PM/PBSK) for the uplink. The uplink data rate is 4 kbit/s; the downlink rate is configurable between 3.5 – 10 kbit/s (the nominal downlink rate is 5 kbit/s). The on-board antenna consists of four λ/4 monopoles with circular polarization, placed in each of the four lateral sides, with a deployment system
The RF communications subsystem, designed and developed at TAS-E (Thales Alenia Space España), is a transceiver that uses the ESA communication protocols such as ECSS E 70 to communicate with the ground segment, and ECSS- E 41 for telecommands and telemetry.
Figure 6: Photo of the transceiver EM (image credit: INTA)
Figure 7: System configuration of the OPTOS nanosatellite (image credit: INTA)
Figure 8: Alternate view of OPTOS with external structure (image credit: INTA)
Launch: The OPTOS nanosatellite was launched on Nov. 21, 2013 as a secondary payload on a Dnepr vehicle from the Dombarovsky (Yasny Cosmodrome) launch site in Russia. The launch provider was ISC Kosmotras. The primary payloads on this flight were DubaiSat-2 of EIAST (300 kg) and the STSat-3 minisatellite of KARI, Korea (~150 kg). 12) 13) 14) 15)
The secondary payloads on this flight were:
• SkySat-1 of Skybox Imaging Inc., Mountain View, CA, USA, a commercial remote sensing microsatellite of ~100 kg.
• WNISat-1 (Weathernews Inc. Satellite-1), a nanosatellite (10 kg) of Axelspace, Tokyo, Japan.
• BRITE-PL-1, a nanosatellite (7 kg) of SRC/PAS (Space Research Center/ Polish Academy of Sciences of Warsaw, Poland.
• AprizeSat-7 and AprizeSat-8, nanosatellites of AprizeSat. AprizeSat-7 and 8 are the ninth and tenth satellites launched as part of the AprizeSat constellation, operated by AprizeSat. The constellation, which was originally named LatinSat, was initially operated by Aprize Argentina; however ownership of the constellation was later transferred to their US parent company AprizeSat. The AprizeSat constellation is used for store-dump communications, and some satellites carry AIS (Automatic Identification System) payloads for Canadian company ExactEarth. The AprizeSat spacecraft were built by SpaceQuest Ltd. Of Fairfax, VA, USA, and each has a mass of 12 kg. 16)
• UniSat-5, a microsatellite of the University of Rome (Universita di Roma “La Sapienza”, Scuola di Ingegneria Aerospaziale). The microsatellite has a mass of 28 kg and a size of 50 cm x 50 cm x 50 cm. When on orbit, UniSat-5 will deploy the following satellites with 2 PEPPODs (Planted Elementary Platform for Picosatellite Orbital Deployer) of GAUSS:
- PEPPOD 1: ICube-1, a CubeSat of PIST (Pakistan Institute of Space Technology), Islamabad, Pakistan; HumSat-D (Humanitarian Satellite Network-Demonstrator), a CubeSat of the University of Vigo, Spain; PUCPSat-1 (Pontificia Universidad Católica del Perú-Satellite), a 1U CubeSat of INRAS (Institute for Radio Astronomy), Lima, Peru; Note: PUCPSat-1 intends to subsequently release a further satellite Pocket-PUCP) when deployed on orbit. 17)
- PEPPOD 2: Dove-4, a 3U CubeSats of Cosmogia Inc., Sunnyvale, CA, USA
MRFOD (Morehead-Roma FemtoSat Orbital Deployer) of MSU (Morehead State University) is a further deployer system on UniSat-5 which will deploy the following femtosats:
- Eagle-1 (BeakerSat), a 1.5U PocketQub, and Eagle-2 ($50SAT) a 2.5U PocketQube, these are two FemtoSats of MSU (Morehead State University) and Kentucky Space; Wren, a FemoSat (1U PocketQube) of StaDoKo UG, Aachen, Germany; and QBSout-1S, a 2.5U PocketQube (400 g) of the University of Maryland testing a finely pointing sun sensor.
• Delfi-n3Xt, a nanosatellite (3.5 kg) of TU Delft (Delft University of Technology), The Netherlands.
• Triton-1 nanosatellite (3U CubeSat) of ISIS-BV, The Netherlands
• CINEMA-2 (KHUSat-1) and CINEMA-3 (KHUSat-2), nanosatellites (4 kg each) developed by KHU (Kyung Hee University), Seoul, Korea for the TRIO-CINEMA constellation. TRIO-CINEMA is a collaboration of UCB (University of California, Berkeley) and KHU.
• GOMX-1, a 2U CubeSat of GomSpace ApS of Aalborg, Denmark
• NEE-02 Krysaor, a CubeSat of EXA (Ecuadorian Civilian Space Agency)
• FUNCube-1, a CubeSat of AMSAT UK.
• HiNCube (Hogskolen i Narvik CubeSat), a CubeSat of NUC (Narvik University College), Narvik, Norway.
• ZACUBE-1 (South Africa CubeSat-1), a 1U CubeSat (1.2 kg) of CPUT (Cape Peninsula University of Technology), Cape Town, South Africa.
• UWE-3, a CubeSat of the University of Würzburg, Germany. Test of an active ADCS for CubeSats.
• First-MOVE (Munich Orbital Verification Experiment), a CubeSat of TUM (Technische Universität München), Germany.
• Velox-P2, a 1U CubeSat of NTU (Nanyang Technological University), Singapore.
• OPTOS (Optical nanosatellite), a 3U CubeSat of INTA (Instituto Nacional de Tecnica Aerospacial), the Spanish Space Agency, Madrid.
• Dove-3, a 3U CubeSats of Cosmogia Inc., Sunnyvale, CA, USA
• CubeBug-2, a 2U amateur radio CubeSat of Argentina (sponsored by the Argentinian Ministry of Science, Technology and Productive Innovation) which will serve as a demonstrator for a new CubeSat platform design.
• BPA-3 (Blok Perspektivnoy Avioniki-3) — or Advanced Avionics Unit-3) of Hartron-Arkos, Ukraine.
Deployment of CubeSats: Use of 9 ISIPODs of ISIS, 3 XPODs of UTIAS/SFL, 2 PEPPODs of GAUSS, and 1 MRFOD of MSU.
Orbit: Sun-synchronous near-circular orbit, altitude = 600 km, inclination = 97.8º, LTDN (Local Time on Descending Node) = 10:30 hours.
Sensor/experiment complement: (APIS, FIBOS, GMR, ODM)
APIS (Athermalized Panchromatic Image Sensor):
The APIS camera features a CMOS detector design. The objective of APIS is to study the degradation of lenses in the space environment - by imaging the same region repeatedly under the same light conditions. The imaging regions of interest are surfaces which provide an image with a constant irradiance (like desert regions).
The optical system uses a refractive objective responsible for focusing the radiation from the scene on the focal plane. The instrument incorporates baffles to shield stray-light radiation. The focal plane employs a two-dimensional array of detectors based on CMOS technology.
Detector size: The CMOS active pixel sensor has an array size of 1.3 mega pixel (1280 x 1024) with a pixel size of 6.7 µm x 6.7 µm. However, the sensor shall use a ROIC (Readout Integrated Circuit) of 640 x 480 pixels in order to boost the readout speed. For that reason, the focal plane size is 4.3 mm x 3.2 mm. The total mass of the APIS device (optics, mechanics and processor) is 120g. Resolution: 273.7 m. Trace H: 175.2 km, Trace V: 131.4 km. Not only snapshot imagery is being done.
Figure 9: Illustration of the APIS instrument (image credit: INTA)
APIS is an athermalized sensor: the payload does not include focusing mechanics in orbit. Defocusing due to thermal environment is corrected by the use of special materials which have been selected as spacers between the lens set and the focal plane. This solution maintains the image quality within a specified temperature range of ±20ºC.
FIBOS (Fiber Bragg Gratings for Optical Sensing):
The goal of the FIBOS imager is to measure temperature by studying the wavelengths of a laser beam travelling across the Bragg gratings. The measurement obtained will be matched with the sensed temperature of a thermistor at the same location. FIBOS uses two sensors attached at different locations, to be able to correlate the perturbations as they may be thermal expansions of materials. 18)
A schematic view of FIBOS is shown in Figure 10:
• The light source unit is a pigtailed tunable laser
• The sensing unit is composed of two FBGs mounted onto a steel support
• The receiver unit is a pigtailed PIN InGaAs photodiode (model EPM605 from JDSU)
• The processing unit integrated in one PCB (Printed Circuit Board) to control the laser, measure the output of the photodetector (PD) and the communications with the distributed on-board DOT (Data Handling Terminal).
Figure 10: FIBOS block diagram (image credit: INTA)
Legend to Figure 10: Iri (electrical current for laser reflectors), Iopi (electrical current that controls the laser optical power), PD (photodetector).
The complete system is 79 mm x 69 mm x 15 mm in size with a mass of less than 120 g and an average power consumption of ~ 1.5 W, according to the requirements of OPTOS. The joints of the source, sensing and receiver units are being made via fusion-splices to create a permanent joint between all the optical fibers. A two-terminal integrated circuit temperature transducer that produces an output current proportional to the absolute temperature (model AD590), is integrated onto the PCB to measure the temperature and compare the data received from FIBOS.
There are two FBGs (Fiber Bragg Gratings) in FIBOS (each one reflects a different wavelength) and they are mounted on a steel mounting with a cantilever to be free of distortion due to the PCB itself. The FIBOS mounting is shown in Figure 11.
Figure 11: One of the FIBOS sensing units with the optics fiber channel to transfer the light from the laser (image credit: INTA)
The processing unit contains the electronics to supply the 5 different electrical currents to the laser, process the signal produced by the receiver unit, follows the synchronization between the laser source and the receiver and contains the electrical interface with the electrical power subsystem and the DOT from OPTOS. The DOT works with a clock basis of several MHz.
The tunable laser, the sensing unit and the PD will be mounted onto the PCB, the PCB will be mechanically integrated into OPTOS as an independent subsystem linked only with the satellite platform via electrical connections.
GMR (Giant Magneto-Resistance) system:
The objective of GMR is to measure the magnetic fluxes produced by Earth’s magnetic field. The GMR device relies on the effect of the change in electrical resistance experienced by a multilayer structure in the presence of a magnetic field. In a GMR, the multilayer structure is composed of magnetic and non-magnetic films. Since the magnetic films are coupled anti-ferromagnetically, it is the magnetization of one film is opposed to the following (it means that the senses of the magnetization of two adjacent films are the contrary - i.e. opposite senses). - If there is not an external magnetic field and a current is applied perpendicular to the magnetization of the layers, then the resistance of the device is very high. In the presence of an external magnetic field perpendicular to the magnetization of the layers, the resistance decreases because the magnetization of the layers rotates following the external field.
Figure 12: Schematic view of the GMR effect (image credit: INTA)
The GMR sensors are physically integrated into chips located on the sides of a cube. The orientation of the measuring axes is that of the X, Y and Z axes of OPTOS. The cube and its coils are shown in Figure 13; the cube size is < 1 cm3.
Figure 13: View of the GMR EBB (Elegant Bread Board) sensor (image credit: INTA)
Aside from the GMR objective, the results of this experiment will also be analyzed in a separate study, as these types of commercial materials have never been tested in the environment of space.
ODM (OPTOS Dose Monitoring) system:
The objective of the ODM device is to measure the radiation environment in space. ODM consists of two on-board units whose measurements will be analyzed and correlated. Each ODM device uses commercial RadFETs (Radiation-sensitive Field Effect Transistors) to absorb the in-situ radiative particles. The total dose will be measured periodically (in minutes) throughout the mission life. In addition, the reference temperature is being measured by a thermistor. These data will be used to correct the measurements.
The ODM main objectives can be summarized in: (Ref. 9)
- To monitor the dose levels in different areas of the satellite
- To correlate the dose values along the mission with the simulated ones obtained with Geant4 using the OPTOS CAD model and standard particle flux models.
- To obtain deviation factor during TID (Total Ionizing Dose) estimation in a space project
- To establish at INTA the expertise to develop (design, integrate, test and data processing) Radiation Effects Engineering sensors.
The low power budget of the spacecraft requires the implementation of an operational mode scheme. It turns out that all onboard services cannot be provided on a continuous basis and in parallel. An efficient power generation can only be provided with a proper spacecraft orientation. Hence, during these maximum power generation periods, the spacecraft can generally not be used for observations or other service functions requiring a considerable amount of energy.
The following modes are defined to optimize the spacecraft services:
• Initial mode: This mode is reached directly after deployment from the launcher. It is being used for warm-up and start-up services of various subsystems.
• Nominal mode: Implies a power consumption mode where all basic subsystems are functioning (OBCom, OBDH, ADCS and EPS).
• Scientific mode: It concerns the services of any of the payloads while maintaining the critical components on. Some payloads shall not be operated simultaneously.
• Observation mode: Requires a pointing maneuver of the spacecraft to the target region to permit imaging by the on-board camera. A sun sensor placed on the –z axis (opposite to the camera) will determine the sun position and maintain a stable attitude during this operation.
• Safe mode: In the event of a spacecraft emergency (loss of attitude, scarce power, etc.) the satellite shall revert into this saving mode, which requires a minimum of power consumption (but maintains a communication capability with the ground).
RF communication periods: These periods represent the highest power demand on the spacecraft. During communication periods, just basic subsystems operations are being maintained (power control, checking, and reduced ADCS operations). The main spacecraft resources will be devoted to TTC (Rx and Tx) functions and subsystem management. No payload operations shall be executed in these periods.
Figure 14: Schematic view of the operational modes of the spacecraft (image credit: INTA)
In February 2009, the OPTOS project passed successfully a CDR (Critical Design Review). Since then, all system designs have been frozen.
The OPTOS ground segment is located at the INTA facilities. The main function of the ground station is to receive information from, or transmit information to the satellite, as it has been designed for its scientific mission. The principal tasks for the earth station will be the following:
• Tracking of the satellite to determine the times of contact
• Telemetry operations to monitor the status of the satellite
• Command operations to control the different functions of the satellite
• Data processing operations to send the scientific data to the user, in the format required.
Figure 15: Overview of the mission segments (image credit: INTA)
Figure 16: Block diagram of the ground system (image credit: INTA)
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.