ISTS 2002-f-15 DEVELOPMENT OF LOW-COST MICRO TETHERED SATELLITE SYSTEM

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1 ISTS 2002-f-15 DEVELOPMENT OF LOW-COST MICRO TETHERED SATELLITE SYSTEM Yuji Sakamoto, Hiroshi Yano, Akiko Nishimine, Motoki Miyazaki, Krishna D. Kumar, Hiroshi Hirayama and Tetsuo Yasaka SSDL, Department of Aeronautics and Astronautics, Kyushu University , Hakozaki, Higashi-ku, Fukuoka , JAPAN ( yuji@aero.kyushu-u.ac.jp) Abstract The development of Tethered Satellite System undertaken by Kyushu University is shown. It consists of 35-kg Mother and 15-kg Daughter. The primary mission is to develop and operate a tethered satellite system, and this requires three important technical missions, which are Tether System, attitude and orbit determination in low cost and self-made satellite bus. For attitude determination, sun sensors and magnetometers are used. And magnetic torquers are used for stabilizing the satellite before stating tether deployment. This paper describes the concept, mission definitions and system design of the tethered satellite system. 1. Introduction Space Systems Dynamics Laboratory (SSDL) in Kyushu University (KU) has been studying the dynamics of a tethered satellite system (TSS). TSS is the generic name of the system which consists of more than two satellites linked by a tether, and it enables broad applications such as an observation in upper atmosphere, orbit transfer system, which are difficult to be conducted by a conventional single satellite. Although the system has been considered for several decades, experiments in space are little. For example, TSS-1 (1992), TSS-1R (1996), SEDS I (1993), SEDS II (1994) and TiPS (1996) were examined in space. On such backgrounds, SSDL started the project "QUEST" since Jan. 1997, which objectives are to develop the TSS for ourselves, operate it in space and acquire the observed data. The results by the project will contribute on the field of TSS researchers. QUEST is a 50-kg class micro satellite. This is expected to be launched as a piggyback Copyright 2002 by the Japan Society for Aeronautical and Space Sciences and ISTS. All rights reserved. 1 satellite, and the size is less than 50 x 50 x 50 cubic cm. It is separated to 35-kg Mother and 15-kg Daughter in orbit, and 2-km tether is deployed. Some universities are participating in the project. Kyushu University (KU) is responsible for the planning of the tether deployment mission, its dynamics analysis and developing Tether System (TDS) which includes Tether Reel Mechanism (TRM) and Subsatellite Ejection Mechanism (SEM). The basic satellite bus system is kindly offered by the other universities and they also have their own missions. The project is still under way. The satellite presented in this paper is different from QUEST but very similar, and some important systems are commonly used in both projects. The codename is "QTEX" which is the satellite undertaken by only KU since Nov The primary mission is the dynamics analysis of a TSS, and the satellite systems are formed only to accomplish this primary objective. The project is simultaneously carried out with QUEST, and both projects can use the common dynamics analysis and hardware of TDS. The all systems of QTEX are designed and fabricated by KU although some technical advices are kindly offered by some institutions such as University Space Education Consortium (UNISEC) which is recently established. SSDL is also participating in the project of CANSAT (since 1998) which is a satellite imitation which is launched in 4-km altitude and dropped off with a parachute. In the project, useful experiences are cultivated to develop an electrical device made by commercial off-the-shelf micro ICs. This experience is very helpful to the development of QTEX. Needless to say, the developing of a satellite costs vast budget, that does not permit the use of commercial off-the-shelf electrical devices because of the failure risk. The concept of QTEX is in the

2 contrary to the conventional thought, and the success will contribute to the future satellite technology. On top of that, the tether deployment mission and the hardware of TDS is very interesting in the present fields of study. The projects of SSDL are summarized in Table. 1 and the aspect and specification of QTEX is shown in Fig Project Summary of QTEX 2.1 Missions The primary objective of QTEX is to develop and operate an experimental tethered satellite system, by which analyzing the dynamics of a TSS will be possible. To achieve the goal, the following three important technical missions are defined, Table 1 Satellite Projects of SSDL, KU QUEST Nov Nov JP/US joint project ASU, KU, SCU Nov KU, SCU, UT, WU CANSAT Nov imitation satellite Many teams from Japan and US launched in 4-km alt. KU participated in launches of by amateur rockets Jul and Aug. 2001, NV QTEX Nov KU's challengeable KU is responsible for project all systems KU: Kyushu University ASU: Arizona State University SCU: Santa Clara University UT: University of Tokyo WU: Washington University 35-kg Mother 50W x 50D x 25H Subsatellite Ejection Mechanism with Springs 20Di x 5H Fig. 1 All length unit is cm 2-km Tether (KEVLAR29) 15-kg Daughter 20W x 20D x 15H Stabilized by Magnetic Torquer Satellite Ejection Mechanism JointedwithRocket 20Di x 5H Aspect as Launched 50W x 50D x 50H Maximum 50 kg Mass Aspect and Specification of QTEX 1. The Daughter jointed with the Mother must be deployed in stable motion. The reliable hardware and stable control algorithm should be studied. 2. To analyze the motion, the attitude and orbit determination system should be developed in low cost. 3. To develop a satellite bus in small budget, self-made satellite bus system should be fabricated by using commercial off-the-shelf devices. By slightly sacrificing the possibility of the success, complex and expensive systems should be excluded. 2.2 Operations The plan of the operations is defined as follows, 1. QTEX is launched as a piggyback satellite in which the Daughter is jointed with the Mother and the TDS is contained in the Mother. The orbit is supposed to be about 800-km sun-synchronous orbit because there is high possibility that QTEX is launched with a scientific satellite adopting this orbit. QTEX does not require the specific orbit basically. 2. After QTEX posed in orbit, the health is carefully checked and the orbit and attitude determination is tried for 2 months. 3. The schedule of the deployment commanded, the tether deployment phase is started. The Daughter is ejected with little initial velocity assisted by springs. 4. The deployment is finished in the tether length of 2 km. This phase is completed in a few hour. 5. The motion of the TSS is observed for at least 4 months. 6. The frequency bands of the communication are ones of amateur radio. Therefore, all the data downlinked can be received to the lovers of amateur radio as well as the researchers of a tether system. 7. After the tether cut by a space debris, two independent satellites keep their operations unless they stop. The operations are summarized in Fig Schedule and Success Level Definition The schedule and the success level are defined in Table 2 and Table 3. 2

3 Ground Station Initial Speed 0.6 m/s Earth 2Km Single Satellite Operation (Health Check, Orbit and Attitude Determination) Tether for 2 Hours TSS Operation (Librational Motion is Observed) Tether Cut by Space Debris Each Satellite Keeps Operation until End of Life Begin of Life 2Months 6Months Table 2 Fig. 2 QTEX Operations Development Schedule of QTEX strength of springs used in the ejection of the Daughter will be required, which causes some troubles. Fortunately, there is much margin mass, and a dummy weight can be loaded in both satellites. There is a possibility that the result of a dynamics analysis requires it. By some simulations, the mass distribution of 35 kg and 15 kg gives a good stable result. The generated power is calculated by supposing the worst case in when only one-side panel is lighted by solar. The generated power is 17 W in the mother and 4.1 W in the daughter after the voltage regulated. In the altitude of 800 km, the daytime is 72 % in the orbital period. Supposing the power consumption is Nov. 01 May. 02 Nov. 02 May. 03 Nov. 03 CONCEPT DESIGN mission definition and system design mathematical model BREADBOARD MODEL design, fabrication and electrical test ENGINEERING MODEL integration of devices and environmental test incl. vibration, thermal and radiation FLIGHT MODEL final fabrication and integration final test LAUNCH wait until opportunity offered ATTITUDE CONTROL & SENSORS POWER Fig. 3 QTEX TETHER DEPLOYMENT SYSTEM QUEST COMMUNICATION STRUCTURE &THERMAL QTEX System Diagram 50% 60% 70% 80% 90% 100% Table 3 Success Levels MINIMUM SUCCESS completion of the development LAUNCH MEDIUM SUCCESS success of communication orbit and attitude determination FULL SUCCESS tether deployment motion determination of TSS discovery of new facts 2.4 System Design The diagram of the system design and the distribution of the mass and power are shown in Fig. 3 and Table 4. The mass ratio between the Mother and the Daughter is derived by the result of simulations in which the stability and probability of successful deployment are researched. When the mass of the Daughter is much smaller than one of the Mother, the tether tension becomes very small which may cause the slack of the tether. Moreover, the stronger Table 4 Mass and Power Distributions Mother Daughter System Device Mass Power Mass Power (kg) (W) (kg) (W) Attitude 3-Axis Magnetic Torquer Control 3-Axis Magnetometers & Sensors 2-Axis Sun Sensors Axis Accelerometers Axis Gyroscopes N/A N/A Electronics (DAC/AMP/RAM) Communication Packet Transceiver (0.5 W, 1200 bps) N/A N/A CW Beacon Transmitter (50 mw) Antennas 0.2 N/A 0.2 N/A Electronics (C&DH/TNC) Power Solar Cells Batteries Electronics (Power Control Unit) Structure & Structure Thermal Satellite Ejection System N/A N/A Cables Thermal Control Tether Tether Reel 1.5 N/A N/A N/A Subsatellite Ejection Mechanism 1.0 N/A N/A N/A System 2-Km Tether 0.2 N/A N/A N/A Electronics (TDS Control Unit) N/A N/A Total Maximum Distribution Margin (= Max. - Total)

4 constant in both of daytime and night, 12 W in the mother and 3 W in the daughter can be used. 3. Mission of Tether The tether deployment is the important mission of QTEX and this is the challenge of a new satellite system. Tether System (TDS) has the following characteristics, 1. The ejection velocity is given by springs. 2. The rotation of the reel is measured. 3. The breaking motor is the only means to control the speed of the rotation. 4. The tether is not retrieved. The direction of the deployment is supposed to be nadir. The example of the failure in the tether deployment is that the reel is stopped before the tether arrives the target length when the Daughter is ejected in the vertical direction to nadir. This is because the gravity assistance is not worked and the minimum tension of the reel consume the energy by the initial velocity. Therefore, magnetic torquers are used to stabilize the attitude before the Daughter is ejected. The required performance of the magnetic torquer can be obtained by dynamics simulations. The control law to achieve the stable deployment, which is no librational motion at the end of the deployment, has been studied in [1] which uses the laws in [2]. This method synchronize the periods of the vibration of the tether length and the in-plane librational motion. The following equations are used in the simulation software, m mm 1 2 e = (1) m + m 1 2 ( ) 2 µ cos 2 S = u θ + θ 3 r (2) ft µ 2 2 l = l Scos 3 φ φ m r (3) e l 3 µ θ = 2( u θ ) φtanφ sin2θ + u (4) 3 l 2 r l S φ = 2 φ sin2φ (5) l 2 Where, m1, m 2 are masses of the mother and the daughter, l, θ, φ are the tether length, in-plane and out-of-plane rotational or librational angle, f is the T tether tension, ruµ,, are the orbital radius, argument of the latitude and gravity constant. These equations are derived by the approximate conditions in which the center of mass is in Keplerian orbit, the tether mass is neglected and the tether length is much smaller than the orbital radius. The stable control law gives the equation of the tether tension in circle orbits as follows, f l l π 1 lref = t+ l t+ l (6) π f 0 sin l f 2 { 3 2 ( 2 l ) 2 3 } = + (7) T me uc lref ucl Where, l ref is the reference length, t is period after the deployment is started, l f is the target length. l0, l 0 are initial length and velocity, u c is the angular velocity in the circle orbit. In some periods just after the daughter is ejected, the reel does not break the rotation. After the minimum tension of the reel reduces the tether speed, it is started to get accelerated by the gravity when this control method begins to be applied. The numerical simulation is examined to study initial conditions giving a good stable deployment. The stable means that the target length is achieved and the in-plane librational motion is under 10-degree amplitude. The parameters are initial in-plane rotational angle and angular velocity. The result is shown in Fig. 4. By this result, backward in-plane initial angle give higher possibility of success deployment than the case of forward initial angle. This implies that the precision of magnetic torquers is not necessary to be very high. If the target angle is chosen to be 30 deg, the pointing error of ± 30 deg are permitted. The result is derived by using the only one control law, therefore, the another control law will be effective in the initial conditions of white area in Fig. 4. Nevertheless, this requires the attitude determination in some degrees of precision. 4. Mission of Attitude and Orbit Determination To observe the motion of a TSS, the attitude and orbit determination system is surely required. If a star tracker and a Satellite Laser Ranging (SLR) were easily available, this problem would be easily solved. However, these system will be high expensive. Therefore, other system should be prepared in low cost. 4

5 initial in-plane angular velocity (deg/sec) Success but Reverse Direction initial in-plane angle (deg) Failure (white area) initial attitude Success (dark area) Fig. 4 Initial Conditions Lead to Success (the case in 2-km tether deployment, 35-kg Mother, 15-kg Daughter) KU has been studying the low-cost orbit determination system in which observed data are Doppler frequencies exclusively. This system is utilized in the orbit determination of QTEX. Referring to the result of Ref. [3], the rotational or libration motion of TSS cannot be measured in this low-cost method. Therefore, to understand the period and amplitude of a TSS, accelerometers and magnetometers are used. The accelerometers and the magnetic torquers requires the correct attitude determination of both satellites. Fortunately, magnetometers and opto sensors, which can be used to the detection of 1-axis sun vector, are available in very low price. GPS was the first idea of low-cost device. However, the commercial GPS has the limit of altitude and it is difficult to use it in space. On top of that, the use of GPS will prevent us from developing a new technology. Consequently, GPS will be the optional choice. 4.1 Orbital Determination by Observed Data of Doppler frequencies Many low-cost micro satellites adopt the amateur frequency bands because the transceivers are easy available and the remodeling cost is not expensive. The budget constructing an amateur radio station is about $5K to $10K including the maintenance. At this station, a packet communication with a satellite is available. Considering the advanced operation, the spectrum analyzing of observed data of Doppler frequencies can be used to determine the orbit. Although this measurement device will cost at least more than $10K, orbit determinations are conducted by only large research institutions so far. To manage the orbit by self ground stations will reduce the operation cost of a satellite. The specifications of Kyushu University Ground Station (KUGS) is shown in Table 5. The theory and covariance estimations have studied in Ref [1][2]. The orbit determination uses the nonlinear sequential batch least squares method in which a state vector at epoch can be updated by the observed data. The spectrum analyzer used in KUGS has the sensitivity of 100 Hz at worst, that is, RMS of observed data will be about 33 Hz. Ref [3] shows the covariance estimation by the RMS of observed frequency. Referring to it, the RMS of the estimated position will be 738 m. In this precision, if the state at epoch is updated at any time, the direction of a satellite is not lost because the beam width of yagi antennas is 30 degrees, which is wide compared with parabola antennas. KUGS is trying the orbit determination by the real observed data of amateur radio satellites. However, the RMS of frequencies after the least squares method is applied does not have constant tendency in which the RMS changes in the range of 1 Hz to 100 Hz. This trial is continuously conducted and the precise characteristic will be clear in a future paper. 4.2 Attitude Determination by Sun Sensors and Magnetometers The specifications of all the sensors of QTEX are shown in Table 6. The demerit of the attitude determination by sun sensors and magnetometers is that it can be available only when the satellite is in daytime. However, this combination will be best choice in terms of cost. The magnetometer HMC1022 is 1-cm micro IC and the opto sensor TSL214 has 64 x 1 linear array in 8 mm. The self-made sun sensor can be fabricated by TSL214 and the box with a slit. Supposing the one IC is used in the range of 60 degrees for detecting 1-axis sun vector, Table 5 Specifications of KUGS Ground Station Location Band Beam Width Spectrum Analyzer Kyushu University Ground Station (KUGS) Fukuoka, Japan 31.21M ALT, E, N 145 MHz, 435 MHz 30 deg. (-3 db) Advantest R3261C Sweep 7 sec/data, RMS 33 Hz 5

6 12 sensors will be necessary in 2-axis measurement. The precision of 1-degree is expected because each sensor has 64 array which detect the strength of sun light. The IRGF geomagnetic model has the precision less than 0.1 degrees in the angle of a magnetic vector. However, the error of the satellite position will affect the error of estimating a magnetic vector. The simulation is conducted in the case of the 1-km RMS of a position error. Fig. 5 shows the RMS of the error angle of a magnetic vector. By the result, the effect of the poor orbit determination will be not large. Magnetic sensors will generate much errors if the proofreading is not appropriate. Therefore, calibration errors will be worst error source. The precision of 1 degree is set to be target of fabrications. 5. Conclusion Kyushu University has several satellite projects, and the concept design of QTEX is introduced in this paper. The primary mission is to develop and operate a tethered satellite system. To achieve this mission, three important technical missions are defined: 1) The hardware and analysis of the Tether Table 6 Attitude Sensors of QTEX Power Device ICs (per unit) Magnetometers HMC1022 (2-Axis) x 3 5V, 5mA Sun Sensors TSL214 (1-Axis) x 12 5V, 4mA Accelerometers ADXL202 (2-Axis) x 3 5V, 0.6mA Gyroscopes ENC-03J (1-Axis) x 6 5V, 5mA Angle Error (deg) System, 2) Low-cost attitude and orbit determination system, 3) Developing satellite bus systems by using commercial off-the-shelf devices. The operation plan, development schedule, success level definitions, mass and power distributions are shown. Moreover, the requirement of the stability before the deployment started was simulated. And the strategy of low-cost attitude and orbit determination are described. The devices of sun sensors, magnetometers and magnet torquers are used in QTEX and now the breadboard model are being test at this time. References [1] Yosuke Nakamura, Tetsuo Yasaka, Hiroshi Hirayama, and Yuji Sakamoto, "Dynamics of Tether from a Small and Cheap Satellite System," 51st International Astronautical Congress, IAF-00-I.2.08, Reo de Janeiro, October 3, [2] Charles C. Rupp, A Tether Tension Control Law for Tethered Subsatellites Deployed along Local Vertical, NASA Technical Memorandum NASA-TM-X-64963, Nov [3] Yuji Sakamoto and Tetsuo Yasaka, "Study of Low-Cost Orbit Determination System for Tethered Satellites," the Ninth International Space Conference of Pacific-basin Societies, Pasadena, CA, November 14-16, [4] Yuji Sakamoto, "Construction of Orbit Determination System Using Low-Cost Ground Station," 22nd International Symposium on Space Technology and Science, 2000-s-10, Morioka, May 30, [5] Daniel Coyle and Henry J. Pernicka, Orbit Determination at a Single Ground Station Using Range Rate Data, The Journal of the Astronautical Sciences, Vol. 49, No. 2, April-June 2001, pp East Longitude (deg) North Latitude (deg) Fig. 5 RMS Angle of Magnetometer Affected by the 1-km RMS Error of the Position 6