11 JSTS Vol. 23, NO. 2
DEVELOPMENT of OICETS ( "KIRARI")
2 2 2 Ichiro Mase1, Kenichi Ikebe1, Naoto Ogura , Koichi Shiratama , Toshihiro Kurii 2 2 3 3 Kazoo Hamada Mitsue Aizono , Akio Yamamoto , Katsuyoshi Arai , 3 3 3 3 Toshihiko Yamawaki , Keizo Nakagawa ,Takashi Jono , Yoshisada Koyama
1 Space Systems Division, NEC Corporation 1-10, Nisshin-cho, Fuchu, Tokyo, 183-8501, Japan. 2 NEC Toshiba Space Systems, Ltd. 1-10, Nisshin-cho, Fuchu, Tokyo, 183-8501, Japan. 3 Japan Aerospace Exploration Agency (JAXA) OICETS Project team 2-1-1 Sengen, Tsukuba City, lbaragi, 305-8505, Japan.
Abstract The Optical Inter-orbit Communications Engineering Test Satellite (OICETS, "Kirari") was developed by Japan Aerospace Exploration Agency (JAXA). The main contractor is NEC Corporation. OICETS was launched into low earth orbit (LEO; altitude of 610km) (Ref.I), leading to the success of two world's first demonstrations in the laser optical inter-orbit communications. The first was two-way communications between OICETS and Advanced Relay and Technology Mission Satellite (ARTEMIS) of the European Space Agency (SEA) in geostationary orbit (GEO) in Dec.2005. The other was communications between OICETS (in LEO) and an optical ground station, in March 2006. The secondary mission of OICETS is the development of a 500 kg class standard satellite bus for J-1 Rocket. This paper presents the design features of OICETS to achieve optical inter-orbit communication and standard satellite bus.
1. Introduction Optical communications are the key technology provided with various advantages comparing with the conventional radio frequencies, including compactness of the equipment, higher data rates, larger communication capacity, and freedom from radio interference. JAXA and NEC have collaboratively been conducting researches on the laser inter-orbit communication technology since 1985. For the in-orbit demonstration of this technology, the Optical Inter-orbit Communications Engineering Test Satellite (OICETS) program was developed, and in 1994, the agreement on the in-orbit communication experiment between OICETS and ARTEMIS was signed with ESA, then the program officially started (Ref.2). When the program started, OICETS was to be launched into circular orbit with a 35° inclination by the J-1 launch vehicle, a compact solid rocket under development in Japan at that time. Developing of a 500kg class standard satellite bus for J-1 launch vehicle was also one of the missions of OICETS. On completion of the protoflight test, however, the OICETS project was frozen for three years because ARTEMIS was not placed in the geo-synchronous orbit due to the malfunction of its launcher, and it took one and half year to settle in the orbit with its own propulsion system. 12 JSTS Vol. 23, NO. 2
In 2004 the OICETS project was resumed, where the launcher was replaced by Dnepr launch vehicle and the orbit was changed to the sun-synchronous orbit (Ref. 3). Since the environment of new launch vehicle and orbit is in the range of the original plan, the original basic design of OICETS was kept intact.
2. Outline ofOICETS As shown in Fig. I, OICETS is a three-axis stabilized satellite with a weight of approximately 570kg. Its main structure is a box type with the dimensions 0.78 x 1.1 x Hl.5 m (Hl.7m including rocket interface ring). When the solar paddles are fully extended, the length of OICETS is about 9.4m in width.
S-band Antenna for LUCE-0 Inter-satellite Link LUCE-0
+Z (Nadia) LUCE-E
Fig. 2 LUCE Configuration Fig. 1 In-orbit Configuration of OICETS
LUCE (Laser Utilizing Communications Equipment) is the mission equipment for laser optical communications experiments with ARTEMIS. As shown in Fig. 2, LUCE is composed of an optical part (LUCE-0) and an electronics part (LUCE-E). LUCE-0, which consists of laser diodes, relay optics, a part of control electronics, and an optical telescope, is a mobile part installed on 2-axis gimbals. LUCE-0 and S-band antennas for inter-satellite link are installed on the top of the satellite for wide visibility of the geo-stationary satellites; ARTEMIS and DRTS. The concept of the inter-orbit laser communication experiment between OICETS and ARTEMIS is shown in Fig.3, and the total communication link for OICETS operation is shown in Fig. 4 13 JSTS Vol. 2:J, NO. 2
I/' I f . I J 1 I J. l ./ .,.. .. ,.. ,S, -----o;er 40.000km ' .k \. ARTEMIS I OPALE termlnel · Geastrioru11yEart. Orbit- OICETSf LUCeterl'l'lil'ld • Law &.arth Orb ft· Fig. 3 Conceptual Sketch of In-orbit Experiments (Ref.1) Olrn< Link·3/ : S-band Telemetry/Command Telemetry _ · · ?' \'1\\ Command \\ I, S-band Telemetry link I- /Command ,, !.l . M 1ssion \\, II Telemr :i!l_I Feeder link KTDS - JAXA GROUND _ HSB STATIONS- USB j OICETS OPERATION CONTROL CENTER ESA REDU STATION OICETS MISSION CONfROL CENTER JAXA Tracking and Control Cent Fig. 4 Total Communication Link for OICETS Operation (Ref .4) 14 JSTS Vol. 23, NO. 2 Communication & Data Handling (C&DH) Micro vibration LUCE Equipment (MVE) Data Recorder Attitude & Orbit Control Switch, Diplexer USB/SSA Central Unit Attitude Earth Transponde /Remote Unit Sensor S-band Antenna Inertia Reference (Gyro) To/From All components Sun Sensor To all component 1------1 Rate Gyro : Thermal ,1 ______(For Spin Phase) _: Distributor Control I I Driver Magnetic Power Structure orque-rod Battery Electronics Controller Corner Cube Electric Power Subsystem fEPSl Paddle Reflector II II = Redundant Fig. 5 Functional Block Diagram ofOICETS 3. System Configuration of OICETS The functional block diagram of OICETS is as shown in Fig. 5. The features are as follows: Micro Vibration Measurement Equipment (MVE) is installed on LUCE mounting area to measure the micro vibration environment that might affect LUCE's highly accurate tracking control. S-band transponders have a USB mode for ground station, SSA mode for inter-satellite link with the data relay satellite, and High speed S band (HSB) mode (BPSK, 1Mbps) mainly for downlink of mission data. Attitude and Orbit Control subsystem (AOCS) utilizes sensors (conical-scan Earth sensor, fine sun sensor, inertial reference unit), and actuators (reaction wheel and magnetic torque-rod). Rate gyro is used for initial spin phase. Reaction Control subsystem (RCS) uses four IN Hydrazine monopropellant thrusters for both orbit and attitude control. Solar paddles generate 1.4kw electric power in total (at BOL, beta angle is 0). The thermal control method is mainly passive control enhanced by local heaters for only critical components (Batteries and RCS). 4. Features of Optical Pointing Inter-orbit communication between LEO and GEO requires extremely precise pointing accuracy due to long distance (about 40,000km) and the sharpness of the laser beam. OICETS achieved accuracy by both LUCE and the spacecraft attitude control system. The optical pointing system and source of 15 JSTS Vol. 23, NO. 2 pointing errors are summarized in Fig. 6. Before detecting the beacon scan beam from ARTEMIS, LUCE points toward the predicted ARTEMIS direction by calculation using the orbit parameters and time. The LUCE pointing system consists of a coarse pointing mechanism (CPM), fine pointing mechanism (FPM) and point ahead mechanism (PAM) as shown in Fig. 6. When the ARTEMIS beam is incident into the field of view (FOV: 0.2° in radius) of the CPM sensor, CPM acts to receive the beam in the center of FOV. After the beam is incident into FOV (0.012° in radius) of FPM sensor, FPM starts tracking activity with CPM. During the beam's round trip between ARTEMIS to OICETS, ARTEMIS proceeds in orbit. This ahead angle is compensated by PAM. The point ahead angle is calculated using orbit parameters and time (Ref. 1). The sources of pointing errors are summarized in Fig. 6. OICETS suppresses pointing errors by the following measures: Open Pointing for Acquisition Tracking LEO Open Pointing PAM LUCE Calculate the target direction PA Angle by orbit ______rrors m Orbit Determination, Time, sensorB=>Sf gimbals :, Fig. 6 LUCE's Optical Pointing System and Pointing Errors 5. Technique for Pointing Optical Ground Station: Inertia Reference Mode (IRM) For laser communication tests between OICETS and the optical ground station (see Fig. 7), OICETS' attitude should be turned upside-down so that LUCE faces the Earth, given that LUCE is installed on an anti-Earth panel (see Fig. 1). For turning upside-down, OICETS employs a simple maneuver; Inertia Reference Mode (IRM) shown in Fig. 8. OICETS attitude is locked by changing the attitude reference from the Earth sensor to inertia reference unit (gyro). After a half orbit revolution, LUCE faces the Earth. 17 JSTS Vol. 23, NO. 2 Laser utilizing .' / OICETS, "Kirari 11 equipment Laser r.ommvnica11on ·•••••••• Satellite control • .... ····•A.. .. JAXA Kirari operation room (Tukuba space center) Fig. 7 Concept of Laser Communication Experiment between OICETS and Optical Ground Station (OGS) (Ref. 5) Upside-down attitude at LST18 opposite side in orbit Eclipse ..../,/ South Pole Fig. 8 Inertia Reference Mode (IRM) for Laser Communication Test with Ground Station 18 JSTS Vol. 23, NO. 2 6. Features of 500 kg Class Standard Satellite Bus for J-1 Rocket In order to establish a standard satellite bus, several innovations were made in OICETS design to minimize size, weight and cost, as summarized in Table 1. Table 1. Features of OICETS for 500kg Class Standard Satellite Bus Item Features Benefits Minimum dead space with easy accessibility Small envelope Mechanical 570kg in same envelope as 390kg satellite: "TAIYO" Configuration Two fully deployable access panels to install/remove Accessibility most components. RCS Both attitude and orbit altitude can be controlled with Light weight only 4 canted thrusters Low cost Mission Data Data rate: I Mbps is available using common S-band Light weight Downlink transmitter for telemetry/command Attitude Earth pointing 3-axis Control Mode Sun pointing Set to any direction (Inertia reference) Flexibility for Orbit Altitude Control range various missions Control :About 170km (Average of perigee and apogee) Thermal Wide variation of sun angle is acceptable Control (beta angle range from -60 to 60°) Battery Nickel Metal Hydride(NiMH) battery High capacity Data Storage Full solid-state data recorder Lightweight Mechanical Configuration has minimum dead space with accessibility to inside components. The upper stage of the J-1 launch vehicle is similar to the Japanese M-3S-2 rocket used for scientific satellites. The fairing size is almost full for scientific satellite "TAIYO" at Im x 1m x 2m and at a weight of 390kg. Retaining the compactness of lm x lm x H2.3m (main body: Hl .Sm), OICETS has a 570kg load including RCS (fuel: 45kg) and redundant bus components which are not in "TAIYO." Two full open access panels enable full access (installation, removal, cabling) to each component as shown in Fig. 9. 19 JSTS Vol. 23, NO. 2 Access Access Panel Panel . Fig. 9 Access Panels m OICETS and 4-thruster Layout Both attitude and orbit altitude can be controlled using only four canted thrusters (see Fig.9) with upside-down maneuver using IRM (see Fig. 8). High rate S-band (HSB) utilizing BPSK mode added to common telemetry/command transmitter provides a 1Mbps mission data downlink without additional mission transmitters. As a standard bus, OICETS targets a wide range of capabilities towards flexibility for future missions. OICETS development realized the following capabilities: Attitude control mode: Almost all major modes are covered. Orbit control: OICETS can compensate for major errors in injection orbit of solid rocket. Beta angle: Thermal control provides a wide variation of beta angles in sun-asynchronous orbit (originally planned for OICETS) The nickel-metal hydride (NiMH) battery and solid state data recorder are common today; however, they were new technologies in the early l 990's when the OICETS project was launched. 7. Conclusion Inter-orbit communications between LEO and GEO requires very precise pointing accuracy due to long distances (approx. 40,000km) and laser beam sharpness. OICETS achieved accuracy by both LUCE and the spacecraft attitude control system. OICETS is featured by the achievement of high accuracy in attitude by the conventional Earth sensor based system. Since the OICETS program includes a mission to develop a SOOkg class standard satellite bus, several technology innovations were introduced to minimize the size, weight, and cost. Acknowledgements The authors would like to thank Sumitomo Heavy Industries (SHI), Fujinon, Canon, NEC Aerospace Systems (NAS) and NEC Engineering (NECE) for their contributions in overcoming tremendous difficulties which led to the success of OICETS and LUCE. 20 JSTS Vol. 23, NO. 2 References 1. T. Jono, M. Toyoda, K. Nakagawa, A. Yamamoto, K. Shiratama, T. Krui, Y. Koyama, "Acquisition, tracking and pointing system of OICETS for free space laser communications", Proceedings of the SPIE 3692, 41-50 (1999). 2. M. Faup, et al., "Experience Gained in the Frame of Silex Program Development and Future Trends", AIAA 16th International Communications Satellite Systems Conference 1996, 782-783 (1996). 3. Toshihiko Yamawaki, Nobuhiro Takahashi, Katsuyoshi Arai et al. "Launch ofKirari (OICETS) from Baikonur," ISTS2006-j-06, 2006.6.5-9 (2006). 4. T. Jono, Y. Takayama, N. Kura, K. Ohinata, Y. Koyama, K. Shiratama, Z. Sodnik, B. Demelenne, A. Bird, K. 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