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3 Development of the Satellite System 3-1 Development of Optical Inter-Orbit Communi- Cations Engineering Test Satellite (OICETS) 3 Development of the Satellite System 3-1 Development of Optical Inter-orbit Communi- cations Engineering Test Satellite (OICETS) YAMAWAKI Toshihiko Optical Inter-orbit Communications Engineering Test Satellite (OICET) “Kirari” is a satellite developed to demonstrate innovative technologies of laser-based optical inter-orbit communi- cations between OICETS and European Space Agency’s (ESA) Advanced Relay and Technolo- gy Mission Satellite (ARTEMIS) and planned to be launched by Japanese J-Ⅰ launch vehicle from Tanegashima Space Center. After its Proto Flight Test has been completed, the plan had to be changed. Finally OICETS was launched by Dnepr launch vehicle from the Baikonur Cosmodrome in the Republic of Ka- zakhstan in August 2005, and succeeded in the world first bi-directional optical communication experiments between two satellites and between low earth orbiting satellite and a ground sta- tion. This paper describes the overview of OICETS and features of its development. Keywords Inter-satellite laser communications, Optical Inter-orbit Communications Engineering Test Satellite (OICETS) 1 Introduction equipments in 1985 and made research model prototypes of the equipments between 1991 Optical communications are expected to and 1993. Based on the results, JAXA investi- promote, in principle, the downsizing and gated if it was possible to conduct an in-orbit weight saving of communication equipments, experiment using ARTEMIS. It then began to and improve their communication speed. Opti- develop Kirari to be launched, using the J-Ⅰ cal communications have great directivity with launch vehicle, to a low earth orbit at an alti- less interference than radio waves, and hence tude of 610 km, with an orbital inclination of have an advantage in transmitting highly- 35°. confidential information. It is therefore ex- The project was suspended after the proto- pected that optical communications can be flight test (PFT) of Kirari, although periodic used for making a compact, light-weight com- tests, tests for life-limited items, and additional munication equipment mounted in a small or tests for installed software were still conducted medium-size earth-orbiting satellite or a deep- to maintain or improve reliability. space exploration spacecraft. When ARTEMIS succeeded in reaching Japan Aerospace Exploration Agency geo-stationary orbit using its propulsion sys- (JAXA) (formerly, National Space Develop- tem, we brought a Laser Utilizing Communi- ment Agency of Japan) began to survey and cations Equipment (LUCE) engineering model examine optical inter-orbit communication (EM) to an ESA optical ground station on Te- YAMAWAKI Toshihiko 13 nerife Island off the coast of Spain in Septem- ber 2003, and performed experiments to check the acquisition and tracking of ARTEMIS in its orbit, as well as communication functions. The project was restarted in 2004 with the launch vehicle switched to a Dnepr rocket and the orbit to a sun-synchronous orbit. Although the situation surrounding the satellite had changed greatly, we did not have to change the basic design of Kirari because it had been de- signed for a broad range of missions with an aim to make it a standard 500 kg-class bus for J-Ⅰ launch vehicles. The satellite was then launched successfully in August 2005 and used for various in-orbit experiments until it was switched off in September 2009. 2 Overview of Kirari development Kirari is a satellite with a launch weight of about 570 kg. The dimensions of the satellite Fig.1 Final appearance of Kirari before launch body are 1.1 m wide × 0.78 m long × 1.5 m high (1.7 m including the rocket connection ring). This rectangular parallelepiped shape the inertial reference frame according to IRU can fit in the φ1.4 fairing of the three-stage J-Ⅰ measurements. When going half way around launch vehicle and also in the fairing of the the earth in “inertia lock” mode, the satellite Dnepr rocket. The satellite deploys two solar points the face where the LUCE-O was array paddles in orbit to produce necessary mounted to the earth, which allowed for the power. When deployed, the distance between communication experiments with the optical both ends of the paddles is 9.36 m. Figure 1 ground station. shows the appearance of the satellite finalized Figure 2 shows a functional block diagram at the launch site. of Kirari, and Table 1 presents the major Kirari usually uses an inertial reference specifications. The mission payload consists of unit (IRU), conical scanning earth sensor the Laser Utilizing Communications Equip- (CES), and a fine sun sensor (FSS) for attitude ment (LUCE) and Micro Vibration measure- detection, and four reaction wheels (RW) for ment Equipment (MVE). LUCE has two main three-axis stabilized attitude control to keep units: One is the optical unit (LUCE-O) which the bottom of the satellite (rocket connection acquires, tracks and points a target satellite side) oriented to the earth. The optical unit of and sends/receives optical signals, and the oth- the Laser Utilizing Communications Equip- er is the electronic circuit unit (LUCE-E) ment (LUCE-O) and S-band inter-orbit com- which sends/receives electric signals to/from munications antennas are mounted on the up- satellite buses and controls the optical unit. per panel of the satellite for clear The electronic unit is installed inside the satel- communications with ARTEMIS and Japan’s lite body. LUCE-O consists of a dual-axis Data Relay Test Satellite (DRTS). gimbal, Cassegrain telescope optical antenna In the communication experiments with with a 26 cm primary mirror, and inner optical the optical ground station, an “inertia lock” unit in which sensors, laser diodes, and relay mode was used to fix the satellite’s attitude in optical systems are installed. MVE is used to 14 Journal of the National Institute of Information and Communications Technology Vol. 59 Nos. 1/2 2012 Communication and Data Handling System（C＆DH） Laser Utilizing Micro vibration S band Semiconductor data recorder Communications measurement antenna Equipment（LUCE） equipment（MVE） Switch，Diplexer USB/SSA Attitude and orbit control system（AOCS） Central unit（CU） Transponder S band /Remote interface unit（RIU） Attitude control Conical scanning antenna electronic unit earth sensor（CES） Reaction wheel Inertia reference All devices （RW） unit（IRU） Fine sun sensor （FSS） Distributor Drive Thermal control （DIST） mechanism system Rate gyro Structure （For spin phase） system 13Ah Ni-MH Power control Solar cell Battery unit（PCU） paddle Instrumentation Paddle system Drive Magnetic system circuit torquer Electric power system（EPS） （PDL） Tank Valve Thruster Reaction control system（RCS） The double squares indicate the unit has a redundant system. Fig.2 Functional block diagram of Kirari measure micro vibrations produced by the weight and cost saving. On the other hand, drive system of a reaction wheel, etc. in Kirari, since the thrusters are all had to be located and to evaluate any influence from vibrations only for increasing the speed of the satellite, on the acquisition and tracking performance of speed reduction control has to be made by us- LUCE. ing the “inertia lock” mode to reverse the atti- The S band transponder of the communi- tude and orient the thrusters in the desired di- cation system has the following modes: USB rection. The tank can contain up to 45 kg of (Unified S Band) mode which is used mainly fuel to compensate for rocket injection error. for house-keeping telemetry and command The thermal control system is mostly passive transmission from/to a ground station, SSA (S system but provides heater control for the bat- band Single Access) mode which is used for tery, reaction control system, and other equip- communication with a data relay satellite, and ment whose allowable temperature ranges are HSB (High-speed S Band) mode which is used critical. mostly for mission data downlinks to a ground In order to make these units and their re- station. The attitude control system not only dundant systems compact, and to maintain has an earth sensor, a fine sun sensor, and an their accessibility, the main body structure is inertia reference unit, but also a rate gyro, designed in such a way that the two side pan- which is used when the solid rocket J-Ⅰ injects els (+X and ‒X panels), which have no solar a spinning satellite into orbit. For the actuator, cell paddles, can be opened. four reaction wheels and three magnetic torqu- ers are used. The reaction control system con- 3 Development of Kirari sists of four sets of thrusters and two tanks. These four sets of thrusters are minimum sets To reduce the development costs of Kirari, to control the attitude and orbit and so lead to a two-step development was employed: the YAMAWAKI Toshihiko 15 Table 1 Major specifi cations of Kirari actions in relation to major problems faced: Item Specifi cations 1) LUCE was developed with original Jap- Weight About 570 kg anese technology, according to the documen- Shape Box shape with two solar cell tation for the interface with the Optical Pay- paddles load for Intersatellite Link Experiments Dimensions Satellite body: 1.1 m × 0.78 m × (OPALE) mounted on ARTEMIS. Due to the (height)1.5 m development schedules of the LUCE and Height with LUCE at top: 2.93 m OPALE, it was planned not to check the inter- Width when paddles are opened: 9.36 m face using actual units on the ground. We Solar cell paddle Power generation: 1220 W or therefore checked the interface in the follow- system higher (EOL, β angle = 0°) ing manner. Highly-effi cient NRS/BSF -The effectiveness of the communication silicon cell sequence, which was important to establish an Power system Astable bus-type distributed optical link, was checked by performing simu- power system 13Ah, Two Ni-MH cells in paral- lations based on a mathematical model created lel by Japan and ESA.
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