TURKSAT-3USAT: A 3U Communication CubeSatA. Rustem ASLAN1), Bulent YAGCI2), Ahmet SOFYALI1), Murat SUER1), Ertan UMIT1), Osman CEYLAN2), Barış TOKTAMIŞ1), A. Selim DURNA1), Erdem AKAY1), Iklim G.AKAY1), and I.T.U. NanoSat Group İbrahim OZ , Şenol 3) 1) GULGONUL3), and TURKSAT NanoSat Group Faculty of Aeronautics and Astronautics, Istanbul Technical University, Istanbul, Turkey 2) Faculty of Electrics and Electronics, Istanbul Technical University, Istanbul, Turkey 3) TURKSAT Inc., Golbasi, Ankara, Turkey
[email protected] As a preliminary effort for Turkey’s national aim of designing native communication satellites, TURKSAT Inc. and Istanbul Technical University(I.T.U.) has combined their experience and capabilities to build Turkey’s first communication CubeSat, TURKSAT-3USAT. It is a 3U CubeSat, with sufficient room for all the systems with full redundancy. Both in house developed systems based on COTS components and readily available commercial CubeSat systems are used together to achieve redundancy. The payload of the 3USAT is a VHF/UHF transponder to be used for voice communication. The satellite main structure is the ISIS 3-Unit CubeSat structure which is a generic, modular satellite structure based upon the CubeSat standard. As a concurrent study, an in house development of a 3U structure is also carried out. The TURKSAT-3USAT contains two on-board computers; two electric power system, and two beacons and modems along with 10 corresponding antenna. Both batteries and super capacitors are employed to store energy. A de-orbiting system is also designed to de-orbit the satellite following its end of life to comply with the current CubeSat standard and United Nations regulations. Key Words: 3U nano-satellite, CubeSat, Voice communication, Passive attitude control 1. Introduction The CubeSats are now considered a disruptive technology revolutionizing the way satellites are built [1]. Most large satellites involve development of custom equipment while CubeSats mostly rely on COTS components due to modest budgets available for their development [2,3,4]. The level of testing is also somewhat reduced since environmental test facilities are expensive and may not be readily available to the developers. The cost and availability of a piggyback launch is currently a major bottleneck, particularly for CubeSat developers from nations without a launch capability. Türksat Inc. Turkey’s national communication satellite operator is planning to design and manufacture Turkey’s future communication satellites, starting with TURKSAT 5X which will follow TURKSAT 4a and b that are contracted to Japan’s Mitsubishi Electric Corporation. This is planned to be a joint effort between related national aerospace companies and research establishments including international collaborations. Currently, there are a number of development efforts towards national earth observation satellites. The very first satellite designed and developed locally in Turkey was ITUpSAT1, a CubeSat prepared by I.T.U. [5]. It was successfully placed in its orbit by India’s PSLV -C14 on September 23, 2009. Today, it is still sending its beacon signals which are regularly tracked by many radio amateurs worldwide [6]. Meanwhile, the radio amateur world is also working towards development of small and affordable satellites particularly ones that include a payload which provides a tracking or communication possibility to the radio amateur world. As a preliminary effort for Turkey’s national aim of designing native communication satellites, TURKSAT Inc.[7] and ITU has combined their experience and capabilities to built Turkey’s first communication CubeSat, TURKSAT-3USAT. It is a 3U CubeSat, with sufficient room for all the systems with full redundancy. Both in house developed systems based on COTS components and readily available commercial CubeSat systems are used together to achieve redundancy. The payload of the 3USAT is a VHF/UHF transponder to be used for voice communication. The satellite main structure is the ISIS 3-Unit CubeSat structure (www.cubesatshop.com) which is a generic, modular satellite structure based upon the CubeSat standard [8]. As a concurrent study, an in house development of a 3U structure is also carried out. The TURKSAT-3USAT contains two on-board computers; two electric power system, and two beacons and modems. Both batteries and super capacitors are employed to store energy. A de-orbiting system is also developed to de-orbit the satellite following its end of life to comply with the current CubeSat standard and United Nations regulations. The antenna deployment mechanism will be used for payload, modem and beacon antennas. The system of antennas, two dual-antenna and two triple-antenna will be located at each four longitudinal-side surfaces of the structure. Antenna deployment will be ensured by melting wire method. Additionally, a spool like system will be located on the side surfaces of the CubeSat by using antenna holders and pins. The power of each mechanism will be controlled by a micro switch. Documentation is playing a key role in a project’s life-cycle. As a quality assurance, the RID (Review Item Discrepancy) forms are created after the PDR and they are being updated by each group after every meeting. These logs enables team to trace of its pace, verify all the steps and result in a more reliable system. TURKSAT-3USAT is planned to be launched by late 2012. 1 2012.The engineering model of it is scheduled to be ready by January. The Transponders consume 2. Overview of the TURKSAT-3USAT As illustrated in Fig. The subsystems of TURKSAT-3USAT Table 1. the transponders are located under the EPS cards to heat the batteries. Therefore. The most suitable electronic card for this region is ITU EPS which includes battery and capacitors on it. In addition. Due to camera viewpoint. Fig.cubesatkit. including extensive qualification testing to be carried out at related ITU laboratories[6].7 Watts by having 40% efficiency. As a result of the thermal analysis. The interior configuration of 3USAT is determined based on following considerations: Geometric constraints of ISIS 3U(www. The distance between two electronic cards in transition region between the two 1-Unit structure is about 50 mm. both COTS systems and in-house development are employed. ACS must be positioned close to mass and geometry center of the satellite. 3USAT subsystems Develop COTS ment Structure İTÜ ISIS3U OBC İTÜ Pumpkin ADC İTÜ Modem İTÜ Astronautical Dev.com) . placed as most top-position card. To provide necessary attitude stabilization and motion damping of TURKSAT 3USAT. 1. 2. The list of development and COTS subsystems is given in Table 1. the thermal analyses show that the middle unit of the cubesat has no any steep change in environmental temperature. Beacon İTÜ Transponder İTÜ TAMSAT Antennae opening İTÜ Battery İTÜ Clyde Space EPS İTÜ Clyde Space Solar Cells 1 panel Clyde Space De Orbiting İTÜ Subsystem The total mass of the 3USAT is about 4kg. 2. both EPS (www. Fig. The TURKSAT-3USAT is developed with maximum possible redundancy. Attitude Control Subsystem (ACS) is a passive magnetic attitude stabilization system consisting of a permanent magnet and magnetic hysteresis rods. TURKSAT-3USAT configuration is composed of modular 1Us. The redundant system architecture of the 3USAT is shown in Fig. camera/second beacon card is 2 Fig. 2. Where possible. 1. Pumpkin OBC (www. The ISIS3U structure is composed of three separate 1-U structure.com) is located as the most bottom card. The Mass budget is given in Table 2. due to the standard PC/104 stack through connector. All subsystems have a backup system with similar architecture.com) cards are identified as top-positioned cards of the first and second units.3. It should be understood that ACS is required to be in the middle of the satellite and the placement of the ACS is made accordingly.clyde-space. TURKSAT-3USAT system architecture .cubesatshop. 79 -20. The results showed that the cubesat will not fail under the worst case static loading conditions where the static analysis was carried out for 15 g vertical and 10 g horizontal accelerations as inertial loads.50 -17.40 oC +85 oC . 244.83 62. Structural Analysis Computational analyses and structural tests of the CubeSat structure are carried out to form a FE Model and to simulate its response to the launch environment.85 33. Table 4.83 -19.e. ISIS3U structure(in the middle) with ITU developed ones The thermal management is one of the most critical phases for keeping the subsystems operational in space environment.76 3 276 Hz 270 Hz 2.38 -21.64 -13.51 -15. no energy requirements and easy implementation.74 -21. 3USAT mass budget Subsystems Mass (gr) Main Structure 580 ADCS 220 Thermal Control 150 Mechanism 225 EPS 1000 Primary Payloads (Two Transponders) 500 Secondary Payload (Camera) 50 CDHS 295 Communication 450 De-Orbiting 330 Margin 200 Total Mass 4000 4. numerical calculations are performed to predict temperatures of the satellite’s outer surface and subsystems (Table 5) which are also compared to analytical solution. all subsystems but the batteries are in operational temperature range and the heaters on Clyde Space EPS are necessary to keep required temperature values during eclipse periods. in 1st Nat.70 37. Comparison of analysis and measurements for ISIS3U 1st Natural Freq. 3 . and Mylar films which prevent heat loss.22 Table 5. between -5oC to 25oC.40 oC +85 oC . the activated heaters are capable to provide the operational temperature for the batteries.29 Having validated the FE model of ISIS 3U structure. thermal systems on cubesats are reviewed and operational temperature ranges of the subsystems are determined.Table 2. 248. Table 4. three different cases are taken into consideration to be able to verify the FE models: Case 1: Main panel FEM modal analysis and testing Case 2: Main panels connected by special link elements FEM modal analysis & testing Case 3: Complete ISIS 3U FEM modal analysis & testing The results of modal analyses & tests agree well as shown in Table 3.34 -20. i. Maximum and minimum temperatures on subsystems Camera & Beacon De-Orbiting ITU Modem Batteries Clyde EPS Transponder 2 ACS ITU OBC ITU EPS Transponder 1 Helium Modem Sensors Pumpkin OBC 20. The results show that.31 -19.42 46.70 oC +70 oC -40 oC +85 oC 3. and 251.56 49.74 59. As a result a passive thermal protection system was decided because of its light mass. there are heaters on Clyde Space EPS to keep operational temperature of the batteries. the first four modes of the 3USAT model were 236.68 35. Following the conceptual design of the thermal system. Table 3. Operational temperature range of subsystems Subsystem Operational Temperature Range Main Structure ADCS Batteries Solar Panel EPS CDHS Antenna Transponder OBC De-Orbiting System Camera . In addition to the passive thermal system.61 43.87 52. Analysis yield a 60 MPa of maximum stress. Moreover. This was demonstrated by an analysis for eclipse periods for EPS with heaters on.42 -10.6. in Case Error (%) Modal Analysis Result Modal Testing Result 1 70 Hz 70 Hz 0. During the structural tests of ISIS 3U.6 Hz.84 26. 3.30 oC +60 oC .00 2 242 Hz 231 Hz 4. Prior to the thermal design.5.6. the model extended to include the entire 3USAT subsystems to predict the maximum stresses and natural frequencies (the natural frequencies must be over 100 Hz according to Cubesat Revision 12 Standards). Thermal Design Fig.ural Freq.30 oC +85 oC -5 oC +25 oC -40 oC +85 oC -40 oC +85 oC .17 -18.54 38.68 -12.74 -14.40 oC +85 oC -40 oC +85 oC .39 42. The passive thermal system on 3U SAT contains Kapton material which covers the outer faces of solar panels. Communication system with transponder has 10 dB link margin.225 MHz ±25 kHz 30 dBm (1W) -120 dBm 150 dB <3 W Yes Monopol (Cu-Be) 8. 145 MHz λ/8 monopole antenna designed. Next stage. 8). VHF uplink signal is up-converted to UHF downlink signal linearly. Frequency up-conversion (mixer stage).200MHz – 435. Main EPS provide power bus 3. It also has 3 MPPT controllers with perturb and observe algorithm. Transponder system specifications are given in Table 6.3V and 5V regulated voltage and a floating battery voltage. is used to amplify signal power level to -40dBm.6 mm by using Ansoft HFSS software tool. filters. Therefore.8x0. The secondary system is being developed in ITU. Engineering model design process continues on a single board sized 88x88mm which is defined in satellite structural layout. Energy is stored using Li-Polymer batteries in 2 series configuration. Durability of the antennas would be a problem because of this length. long antenna structure complicates the deployment by a wire-melting system. but they have much longer lifetime and temperature characteristics.250MHz downlink frequency band. 3. Main EPS has 3 MPPT charge controller one for each axis. Two lithium polymer 4 . All sub-systems have their own antennas for safety of communication.9).5 cm <250gr 2. transponder and beacon (Table 7). System has 150dB gain with 60dB of the total gain being adjustable. Receiver part includes a very low noise accorded amplifier and narrow band (50 kHz) image filter. thus an impedance matching network is necessary to deliver maximum power. This length is very long so this may cause problems on satellite surface for placing antennas and deployment system. gain control unit protects power amplifier input and output power can be changed from ground station remotely. Additionally.940MHz – 145. Otherwise.8x8. Energy is stored in Supercapacitors. MMCX connector and connection points to hold antenna stable on the surface of the satellite are modelled by using Ansoft HFSS simulation tool considering the actual structure characteristics. gain control unit protects power amplifier input and output power can be changed from ground station remotely. gain control and transmitter parts. Electric Power System TURKSAT-3USAT have two redundant EPS’s. 5V DC MSP430F5438 Fig. furthermore satellite is the first communication satellite which designed in Turkiye. Table 6.8V. Antenna System Satellite communication system includes 3 main devices (sub-systems) with their backups which are modem.990MHz uplink and 435. filter and several gain block. Transponder consists of three main sub-units: Receiver. Subsystems and their antennas UHF Antenna VHF Antenna (435 MHz) Transponder Modem Beacon TOTAL 2 2 2 6 (145 MHz) 2 2 4 4 4 2 10 System Antennas The signal is amplified to 1W before transmitting and this block’s gain is controlled by automatic gain control system to fix the output power to 1W (30 dBm) to keep power consumption stable. Last stage includes a power amplifier which is approximately 50% efficiency with 32dB gain control ability. Transponder consists of amplifiers. The signal is amplified to 1W before transmitting and this block’s gain is controlled by automatic gain control system to fix the output power to 1W (30 dBm) to keep power consumption stable. 8. TURKSAT – 3USAT has 145. 8. Electromagnetic simulations are used to obtain best efficiency for 10 antennas to find locations on the satellite taking into account mechanical constraints. Consequently. The main unit is 3U EPS from Clyde Space. Transponder desktop model 6. Table 7.6 – j251. gain control block includes a variable gain amplifier which adjusts the power between -40 dBm and 5 dBm. Transponder Transponder is the main payload to provide two way radio communications. λ/8 monopole antenna does not have 50Ω impedance (Z = 40. Female 50 Ω MMCX connector is used in this project between the devices and antennas.3 mm and 3. On a desktop model electrical and radio properties were tested (Fig.965 MHz Downlink Center Frequency Bandwidth Output Power Receive Sensitivity Gain Total power consumption Automatic Gain Control Antennas PCB sizes Mass Supply Voltage Microcontroller 435. As a result. 145 MHz and 435 MHz monopole antennas thickness and width are determined as 0. Theoretically. it is impossible to locate 10 antennas without any collision and magnetic interference at each surface. Additionally. Batteries have a heating system for cold eclipses. Two surfaces on the satellite are reserved for antennas. 1W output power. system has 150 dB gain.3V. Transponder system specification Uplink Center Frequency 145. monopole λ/4 antenna length for 145 MHz (VHF) is 574 mm. mixers and AGC (Automatic Gain Control) stages. Transponder is the payload of the satellite.5. Supercapacitors provide less Wh per kg. nearly 35-40% efficiency (Table 6). Furthermore. gravity gradient and solar radiation pressure are simulated for Sun-synchronous. and the transferred measurement data will be processed on ground to determine the attitude states corresponding to angular motion measurement sets. To counteract the undesired torques due to the residual magnetic moment. the magnets onboard are not allowed to have elongation such as four or five.3 1 (min) 35 35 15 5 10 Energy (Whr) 1. which should be about 500mA in orbit. angular motion will be only sensed by the 9 degrees of freedom sensor unit of the satellite. Position of the top layer is adjusted for the camera module. If the disturbance torques emerging from aerodynamic drag. Table 9.6333 0. One of the solar panels will be produced in ITU.1666 2.3 Wh. Every satellite has a residual magnetic dipole moment value that emerges from inductions in electrical circuits onboard.0583 0. 3. altitude [km] Td [N/m²] βexp [deg] 20 10 5 Necessary dipole moment onboard 600 800 8. If that angle is symbolized by β then. The resolution is VGA and the output is JPEG picture format. Energy Requirement (worst case) Transponder OBC Beacon Attitude Det. Attitude Determination and Control System Fig. but after 2 min of illumination the panels does not show too much warming. In other words. mpm = Td/(Bminβexp) can be written. Camera The mission has no precise pointing requirement. On the structure. 5 Due to the restricted volume for the attitude control card in the designed satellite.. There will be temperature sensors and reverse protection diodes on the solar panel. The necessary mpm values are given in Table 9. Later on. that increases the angle between the magnet dipole moment (mpm) vector and the geomagnetic field vector. hysteresis rods produce a damping torque thanks to the energy dissipation property of soft ferromagnetic materials. a hole is put to stick the lens of the camera module.supercapacitors have implemented in series configuration to store energy during eclipse. a LED based solar simulator is tested. power consumption and simplicity. the cells will be covered with a cover glass from Qioptiq [10]. Differently. The operating temperature is from -25oC to 60oC. 9. Standard operating voltage is between 2. The components of such a passive control system are “permanent magnets” and “hysteresis rods”. Therefore.574x10-2 C329 UART camera module is chosen.Thus AlNiCo-8 magnets are chosen.619x10-2 Am². due to its size. Sİnce no active control will be employed. The cell short circuit current was 430mA. It is tested with regular ground use silicon based solar panels. A demonstration model is built and tested for the developed EPS system. the resistance to demagnetization.35 mm and height of 9. therefore no heating or cooling system is required.08x10-1. which are hard ferromagnetic materials. no onboard attitude estimation is required.806x10-8 5.817x10 3.794x10-2 3. is calculated to be 2.025 0. Td. which is good for long term testing.595x10-8 mpm [Am²] 9.5 0.8V for the chosen supercapacitor [9].12. Table 8 shows that required energy for the worst case is about 2 Whr. .107x10-3 1. Fig.619x10 -2 -2 9. permanent magnets. It is complicated to estimate its value therefore it will be taken into account as a disturbance component. Table 8. An UART interface is implemented to control the camera and pass the picture information to OBC or Modem. which can be manufactured with elongation values of two or less thanks to their higher coercivity.8 0. which are typical minimum values for available AlNiCo-5 magnets. B. thus the satellite will be controlled passively by onboard magnetic moment production.125 0. which is within the limits of thermal simulation results. Permanent magnets that pass through geomagnetic field in orbit produce a magnetic restoring torque that is proportional to their total magnetic dipole moment value. circular orbits with altitudes of 600 and 800 km by using realistic mathematical models that take the momentary attitude state of the satellite into account. exchange the angular momentum of the satellite while hysteresis rods reduce it by converting a part of the kinetic energy of the angular motion to heat and consequently reducing the angular velocity.008 10.1 0. therefore the storage can supply the entire system even for the worst case. It was seen that MPPT charge and regulators work without any observed problem.223x10-3 1. The restoring torque counteracts the components of the environmental disturbance torques. If magnets of AlNiCo-8 with diameter of 6. The LED solution does not provide enough radiation to simulate orbital sun energy. A quick demonstration panel is built from a 2 sided PCB. Due to our previous experience with solar simulators. Demonstration of developed EPS Solar panels will be provided mostly by Clyde Space. Two Solar panels are attached on panel. where Bmin is the minimum magnitude of Earth’s magnetic field encountered in orbit and βexp is the minimum expected value of β [ 11]. Fig. 12. Therefore stored energy will be = 2. Modem TOTAL Power budget requirements Power Time (W) 2.2V and 3. which is 3.53 mm are preferred then the number necessary to reach the maximum mpm value in Table 9. A 100W LED with reflectors and optics is used to light the cells. S.itu. Vol 43. et. Ovchinnikov. As the result of the hysteresis tests on the developed test stand. For 3USAT.org/images/developers/cds_rev12. the actual values of these three parameters have to be found.”. Vol 64. Istanbul.C. The layout of the components of the passive magnetic attitude control system on the PCB should match the following criteria for the best performance [17]: Hysteresis rods are to be placed in the plane perpendicular to the permanent magnet dipole axis in order to obtain damping torque as required. 9. Acta Astronautica. M. “CubeSat Design Specification. Vol 46.qioptiq.: “Passive magnetic attitude control system for the Munin nanosatellite”. I. V. Ovchinnikov. D. the coercivity. V. L. 2009. k and V are respectively the dimensionless form factor and volume of the rod. 2007. Tsukuba. Pen’kov.0578 T. pp. J. 27th ISTS (International Symposium on Space Technology and Science).. Yu: “Two modes of the angular motion of artificial satellites. 27th ISTS (International Symposium on Space Technology and Science). pp. F. Battagliere. SSC10-IV-4.13 A/m. pp. Ovchinnikov. et. 803Ovchinnikov. M. Clark. Battagliere. Issues 2-6.tr. ISTS 2009-f-10.: “Emerging Opportunities for Low-Cost Small Satellites in Civil and Commercial Space”. Keldysh Institute of Applied Mathematics. 10. Building and Experimental Results of a Facility to Test Hysteresis Rod Parameters”. it is planned to place three hysteresis rods on each of the rod axes. 8. Keldysh Institute of Applied Mathematics. the calculated necessary magnet amount should be multiplied by a safety factor. Acknowledgments The project is fully supported by TURKSAT Inc. ISTS 2009-d-05. Moscow. 31May-4June. M.: “Attitude control system for the first Swedish satellite “MUNIN”. Hysteresis rods are to be placed orthogonally to each other to obtain the best performance...: Methods to obtain the principal parameters of simple attitude control systems for small satellites. T.jp/productdata/sheet/lic2540r3r8207_e. N.5πMr/Ms)/Hc] M=(2Ms/π)tan-1[k(H±Hc)] mhys = MnV where H is the magnetic field strength along the axis of the hysteresis rod. 319-326. of Turkey. 5-7 June 2008. Grazziani. 4. 2011. the magnetization characteristics of the specimen can be measured by observing the inducted voltage in a second coil around the solenoid according to the “Newman-Faraday-Lenz induction law”. V. In Ref [11]. F. Portugal. and Ms = 0. 11.. M.: “ITUpSAT1: Getting Ready for Launch”. Turkey. Hc.unpredictable changes in space environment.. The present study will present what can be achieved with the developed systems for the communication performance.turksat. M. October 2011. and Ms are respectively the magnetization. I. O. The selected safety factor has to be judged through simulations to decide on its ultimate value. and Barabash.. Acta Astronautica”.: “About Nano-JASMINE Satellite System and Project Status”. and Zelli. Norberg. 12. Mr = 0. Kurtuluş. 2000. and McKee. The 4S Symposium. S. pp. 701-708. D.edu.13. and Penkov. 1998.co. M. Issue 2. V. 24th Annual AIAA/USU Conference on Small Satellites.: “Passive magnetic attitude stabilization of the UNISAT-4 microsatellite. http://www. k = tan[(0. http://cubesat.pdf http://www.. Since the characteristic values Hc.: “Medium Resolution Multispectral Earth Observation from a 5Kg Nano-satellite”. 597-605.yuden. Acta Astronautica”. Moscow. Y. which corresponds to the usage of 8 magnets. Tsukuba. M.14. 142-156. International Workshop on Small Satellites: New Missions and New Technologies. Russian Academy of Sciences. Jul 2009. Kupriyanova. and Ms that determine the hysteresis cycle of the rod differ from the catalogue values for HyMu80.com/. October 2011. 6 . References 1. 6. the measured values of the ferromagnetic parameters are Hc = 52. M. 16. a factor of 16 is acceptable. no: 65. 15.. Vol 54.: “Attitude Determination and Control System in Pico-satellite for Remote-sensing and Innovative Space Missions (PRISM)”. The damping magnetic moment is produced by hysteresis rods according to the relations. Henry. 17. Vol 40. If the rod is put into a solenoid and a magnetic field is created inside the solenoid. If the hysteresis rods are in the equatorial plane of the magnets then there is theoretically no interaction between magnets and rods even if they are in close proximity.15].tr. and the saturation magnetization of the rod. 5.. 13. Issue 5.022 T. and Penkov. F. 2.. N. Jul 2009. 2010. M. V. Mr.: “Preliminary Experimental Results of a Facility to Test Hysteresis Rod Parameters: Effect of the Magnetic Field of a Permanent Magnet”. Kupriyanova. Greenland. Mr. Santoni. which says that an alternating current passing through the solenoid induces an electromotor force along the induction (second) coil. It can be considered as a preliminary effort for Turkey’s national aim of designing native communication satellites considering Turkish national satellite operator TURKSAT’s plan to produce the first GEO communication satellite in 2015. Graziani. the resonance issue between the natural frequencies of the satellite and the frequencies of perturbations arising from non-uniform rotation of the local geomagnetic field vector due to the satellite’s orbital motion has to be investigated [12. al. For an analytically sophisticated decision. Sako. Yu. and possible collisions with debris that are not fatal. http:// http://www. The parallel distances between the hysteresis rods have to be at least 0. 2002. M.com. 7. and n and m are respectively the number of rods and the induced magnetic dipole moment along each principal axis [16]. the remanence. 14. Foust. which is the material of the manufactured rods and is an Iron-Nickel-Molybdenum alloy with 80% Nickel composition. pp.: “Design. S. 9. Contributions of Hasan Körük and Kenan Yüce Sanliturk to the structure development are gratefully acknowledged. Y. Rev..al. 3. Russian Academy of Sciences.pdf QIOPTIC website. Inamori. Madeira.12. ITU.3 times their length so that they can act individually. N. that factor is taken as 30 for the microsatellite UNISAT-4. http://usl. and Ovchinnikov. Journal of Applied Mathematics in Mechanics”. Restricted by the available space and inspired by previous missions. C. no:66. 2000. Issues 11-12. Cosmic Research. Conclusions A nano-satellite with full redundancy is developed for voice communication at low earth orbit. 2007. Ovchinnikov. L. 2010. 792-.