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ChubuSat-1 (Japanese Microsatellite / Kinshachi)

Overview    Spacecraft    Launch   Sensor Complement   References

ChubuSat-1 is a Japanese microsatellite technology demonstration mission, jointly developed by a consortium of universities (Nagoya University and Daido University) and aerospace companies in the Chubu region of central Japan, which is the core region of Japan's aerospace industries. ChubuSat-1 is developed to expand the satellite business by significantly reducing the cost of spacecraft and applications, and also to demonstrate the advanced technology level of the small aerospace businesses in this region. 1) 2) 3) 4) 5) 6)

This program is the result of a strong industry-academia-government collaboration. Funding is provided by JSPS (Japan Society for the Promotion of Science) in the FIRST (Funding Program for World-Leading Innovative R&D on Science and Technology) program , initiated by CSTP (Council for Science and Technology Policy) of Cabinet Office, Government of Japan.

The four satellites to be developed in this program by Japanese institutions are:

• Hodoyoshi-1 :The University of Tokyo and NESTRA (Next Generation Space System Technology Association)

• ChubuSat-1 : Nagoya University and Daido University

• Tsubame : Tokyo Institute of Technology, Tokyo University of Science and JAXA

• QSAT-EOS : Kyushu University.

The nickname of the ChubuSat-1 mission is "Kinshachi", meaning golden grampus, which is the roof monument of the Nagoya Castle, the symbol of the Chubu region.

The ChubuSat-1 mission has the following objectives:

1) Observation of Earth's surface

2) Observation of space debris

3) Relay of amateur communication data (provision of a message transfer service to worldwide amateur radio users).

ChubuSat-1 has an optical camera and an infrared camera to observe Earth's surface. The resolution of the optical camera is approximately 10 m. The infrared camera is sensitive in the wavelength from 7.5 to 13.5 µm and enables observation of the ground temperature profiles.


Figure 1: Artist's view of ChubuSat-1 on orbit (image credit: ChubuSat consortium)

Development team: Nagoya University, Daido University, and MASTT (Meiyu Aerospace Support Technology Team), collaborative aerospace companies in the Chubu region of Japan, organized the ChubuSat-1 project team (or ChubuSat consortium). In this setup, Nagoya and Daido Universities are in charge of the ChubuSat-1 satellite system design, procurement of satellite components, satellite assembly, system testing and evaluation. Nagoya University also takes charge of the construction of the ground station systems and processing and evaluation of the observed imagery. MASTT is in charge of manufacturing most of the satellite mechanical or electrical parts, except for the purchased components.



The microsatellite is almost a cube of size: 58 cm x 55 cm x 50 cm and a mass of ~ 50 kg. The design life is 6 months.

Structure: The ChubuSat-1 design employs a standardized, modularized, general purpose and flexible satellite system architecture. The main structure consists of sandwich panels of aluminum skin and aluminum honeycomb core which has light mass and high stiffness. The light panels, whose total mass is < 10 kg, meet both the stiffness requirements from the rocket and the mass requirement from the system design. The structure of ChubuSat-1 has been developed from the STM (Structural and Thermal test Model) which was tested by a modal survey and vibration test with qualification test level of the Dnepr launch vehicle.


Figure 2: Illustration of the ChubuSat-1 microsatellite (image credit: ChubuSat consortium)

TCS (Thermal Control Subsystem): TCS is based on passive control using MLI (Multilayer Insulation) and radiation film of silver Teflon. Use of heaters and temperature sensors.

ADCS (Attitude Determination and Control Subsystem): The spacecraft is 3-axis stabilized using the zero-momentum bias technique. The attitude is sensed with a star sensor, sun sensors, and 3-axis magnetometer, 3 FOG (Fiber Optical Gyroscopes) provide the rate information. Actuation is provided by reaction wheels and magnetic torquers. During the initial phase after deployment from the launch vehicle, or during a safe-hold phase, two-axis sun-pointing attitude control is used. The magnetic torquers are used used for unloading of wheels and for the rate dumping of the separation tumbling motion.

The following attitude control processing functions are implemented by the OBC software:

- Rate control

- Sun-pointing 2-axis control

- 3-axis control to commanded quaternion

- Attitude maneuver path generation

- Wheel unloading

- Attitude FDIR (Failure Detection, Isolation and Recovery ).

The ADCS pointing requirements call for a 3-axis pointing accuracy of < 0.8o and for a 2-axis sun pointing accuracy of < 5o.


Figure 3: Block diagram of the ChubuSat-1 microsatellite (image credit: ChubuSat consortium)

EPS (Electrical Power Subsystem): The EPS consists of a PCU (Power Control Unit), a NiMH secondary battery, and three solar array panels. As the center of the power control system, the PCU has the following functions:

1) Regulation of solar array power: regulated by partial shunt control.

2) Secondary battery charging/discharging: switching constant current charge mode and trickle charge mode based on battery voltage.

3) Battery temperature monitor: this function stops the charging when the cell temperature goes beyond the threshold.

4) Power-on and shut-down: the power circuit is activated with two flight pins and two separation detect switches. When ONLY both switches are switched on, the power supply will be started. On the other hand, in case of the mission termination, the solar array power lines are cut off by discrete command from the ground station, and satellite will be powered down when the rest power of the battery is used up.

5) Power distribution: generate +5 V, ±15 V, and ±12 V of power from the +28 V non-regulated bus voltage, and distribute this power to each component.

6) Component power supply ON/OFF: switch by a command from the OBC.

7) Housekeeping telemetry output: respond to the power control status from an OBC request.




PCU (Power Control Unit)

Primary bus voltage

23 - 36 V

Maximum handling power

85 W

Secondary battery (NiMH)

Number of cells

21 / string x 5 strings


1.9 Ah / string x 5 strings

Discharge voltage

23 – 29 V @ 20% DOD

Charge/discharge cycle

>10,000 @20% DOD

Solar array panels


> 100 W

Solar cells

GaAs triple junction

Number of solar cells

20 / string x 3 strings/panel x 3 panels


29.1% @28oC

Table 1: Specification of the EPS

OBC (On-Board Computer): The OBC consists of three module types, the core module with the CPU and memory, the extended interface module which has interfaces to the other components, and the PCU Power Control Module). The core modules and extended interface modules are connected with SpaceWire interfaces. This enables future extensions of the OBC functions by the addition of a core module or an extended interface module.

The ChubuSat-1 OBC has two core modules and three extended interface modules. It should be noted that a SOI-SOC (Silicon On Insulator, System On Chip) chip is part of the core module; it reduces power consumption, and endures a high temperature and a high radiation environment (Figure 4).


Figure 4: Photo of the OBC flight model (image credit: ChubuSat consortium)

The OBC can be modified easily by changing and modifying interface circuits and software. This gives us flexibility of adopting future satellite/components. Moreover, the OBC is an integrated component of the AOCU (Attitude and Orbital Control Unit), DHU (Data Handling Unit), DR (Data Recorder), and the CMD (Command Decoder). This integration reduces testing cost, wire harness, and assembly workload.

The OBC has the following functions:

• Boot control: self-check, memory initialization, and watchdog timer clear.

• Operation control: on-board software control, external interface control, satellite time control

• Attitude control: calculate torque command to each actuator using attitude output from each sensor.

• Data processing: command decoding, telemetry gathering, editing and encoding, and communication between peripheral components.

• Mission component control.

• Prevention of overcurrent failure propagation: when overcurrent is detected, the failure module is shut down so that overcurrent would not propagate to other modules anymore.

• Failure detection, isolation and the reconfiguration (FDIR) function.

• Reprogramming capability: The ChubuSat-1 OBC is reprogrammable on orbit.

Processing speed, data bus width

50 MIPS, 32 bit

Memory storage / core module

1 MByte (EEPROM)
2 MByte (SRAM)
64 MByte (SDRAM)
512 MByte (Flush ROM)

Radiation resistance

TID:>20 krad, SEL:> 30 MeV/mg/cm2

Power consumption

< 16 W

Instrument size, mass

270 mm x 205 mm x 130 mm, 5 kg

Table 2: Specification of the OBC

RF communications: The communication subsystem uses amateur radio frequencies. The data rate is 1.2 kbit/s in uplink (VHF) and 9.6 kbit/s in the downlink (UHF). The communication subsystem consists of two amateur radio transceivers (A-TRX), a transmitter switch, a transmitting antenna switch, two receiving antennas and two transmitting antennas.


Figure 5: Block diagram of the communication subsystem (image credit: ChubuSat consortium)


Figure 6: Photo of the A-TRX (Amateur Radio Transceiver), image credit: ChubuSat consortium

ChubuSat-1 features a receiving antenna hold and release mechanism (Figure 7). In the launch configuration, the receiving antennas are folded and held by this mechanism with nylon cable. A Nichrome wire is wound around the nylon cable and connected to the PCU. After satellite separation, the PCU heats the Nichrome wire resulting in a cut of the nylon cable, and the folded antennas will be released and deployed by a spring torque. The nylon cable holding the antenna is multiplied to meet the range safety requirements.


Figure 7: Receiving antenna hold and release mechanism of ChubuSat-1 (image credit: ChubuSat consortium)


Figure 8: Overview of the communication subsystem (image credit: ChubuSat consortium)

ChubuSat-1 will provide a global message relay service using the onboard transceiver of the amateur radio band.


Launch: The ChubuSat-1 microsatellite was launched as a secondary payload on November 6, 2014 (07:35:49 UTC) on a Dnepr-1 vehicle from the Yasny Cosmodrome, Russia. The primary payload on this flight was the ASNARO minisatellite of USEF, Japan.
The launch provider was ISC Kosmotras. The launch was executed by the Russian Strategic Rocket Forces of the Russian Ministry of Defense with the support of the Russian, Ukrainian and Kazakhstan organizations, which are part of the ISC Kosmotras industrial team. 7)

Orbit: Sun-synchronous orbit, altitude of 504 km, inclination = 97.4o, LTDN (Local Time on Descending Node) = 11:00 hours.

The secondary payloads on this mission were: 8)

• QSat-EOS, a microsatellite (49 kg) Kyushu University (KU), Fukuoka, Japan.

• Hodoyoshi-1, a microsatellite (60 kg) of the University of Tokyo and NESTRA (Next Generation Space System Technology Association)

• Tsubame, a microsatellite (49 kg) of Tokyo Institute of Technology, Tokyo University of Science and JAXA

• ChubuSat-1, a microsatellite (50 kg) of Nagoya University and Daido University, Japan.



Sensor complement: (VIS Camera, TIR Camera)

ChubuSat-1 features an optical and an infrared camera to observe Earth's surface. The resolution of the optical camera is ~10 m. The infrared camera is sensitive in the wavelength region of 7.5-13.5 µm and enables observations of the ground temperature profile. The project also plans to observe space debris using the infrared camera.

VIS (Visible) Camera
(developed by Tokyo University of Science)

Spectral range

0.4-0.8 µm

Detector, size

CMOS, 2048 x 1536 pixels, pixel pitch = 3.5 µm

FOV (Field of View)

2.15o x 1.61o

Spatial resolution

10 m

Storage capacity

2 GB


180 mm x 75 mm x 75 mm, mass = 0.6 kg

Electronic box

110 mm x 60 mm x 40 mm, mass = 0.3 kg


< 5 W (5 V input)

Operating temperature


TIR (Thermal Infrared) Camera
(COTS product)


Spectral range

7.5 to 13.5 µm

Detector, size

Uncooled bolometer, 320 x 240 pixels, pixel pitch = 25 µm


4.6o x 3.7o

Spatial resolution

130 m


< 1 W (5 V input)

LVDS/CMOS interface


Device size, mass

80 mm x 80 mm x 150 mm, 0.5 kg

Table 3: Specification of the camera performance parameters


Figure 9: Photo of the VIS camera (image credit: ChubuSat consortium)


Figure 10: Photo of the TIR camera (image credit: ChubuSat consortium)


1) Yasutaka Narusawa, Hiroyasu Tajima, Masanobu Mizoguchi, Isao Matsubara, Hirotaka Kosaka, "Development of ChubuSat-1 Satellite," Proceedings of the UN/Japan Workshop and The 4th Nanosatellite Symposium (NSS), Nagoya, Japan, Oct. 10-13, 2012, paper: NSS-04-0315, URL:

2) "Development of ChubuSat-1 Small Satellite," April 9, 2013, URL:

3) "Objectives and Overview of ChubuSat Instrument Development Projects," URL:


5) Hiroyasu Tajima, "ChubuSat Project," Nagoya University, March 11, 2013, URL:

6) Seiji Yoshimoto, et al., "Environment Monitoring of Fukushima and Chernobyl Areas using a Constellation of Earth Observation Microsatellites," University of Tokyo, Nov. 20, 2013, URL:

7) "Dnepr Launch of ASNARO and 4 piggyback microsatellites," ISC Kosmotras, Nov. 6, 2014, URL:

8) "Japanese Microsatellites Launched," AMSAT, Nov. 6, 2014, URL:

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 (

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