Rising-2 is a cooperative microsatellite project of Tohoku University (Sendai) and Hokkaido University, Sapporo, Japan. The primary objective of the mission is Earth observation with a resolution of ~ 5 m. In particular, high-resolution cumulonimbus scenes will be observed using the LCTF (Liquid Tunable Multispectral Filter) technique. The secondary objective is the observation of sprite phenomena in the upper atmosphere. Sprites are so-called TLEs (Transient Luminous Events), which are rather frequent natural phenomena induced by lightning discharges. 1) 2) 3) 4) 5) 6) 7) 8)
The Rising-2 project started in July 2009 and inherited the experiences gained in the design and development of the SpriteSat (Rising-1) spacecraft which was launched on January 23, 2009. The Rising-2 microsatellite is expected to be completed in the summer of 2013.
Figure 1: Two views of the Rising-2 microsatellite (image credit: Rising-2 partners)
The spacecraft structure is cubical with a side length of 50 cm and a launch mass of ~42 kg. The central pillar configuration is inherited from SpriteSat, using an aluminum alloy material for the central pillar and the side panels. The spacecraft is 3-axis stabilized.
ADCS (Attitude Determination and Control Subsystem): The ADCS uses star and sun sensors, gyros and magnetometers for attitude sensing (Honeywell HMC2003). Actuation is provided by a reaction wheel assembly and 3 magnetorquers (MTQs). The observation requirements call for a spacecraft pointing accuracy of < 0.1º and an angular velocity accuracy of 0.02º/s. 9) 10) 11) 12) 13)
The ADCS offers two observation modes:
• Fine pointing control mode: This is the observation mode which is used for ~ 15 minutes in the sunlit phase of the orbit and for about 15 minutes in the ecliptic phase of the orbit (use of star sensor and reaction wheels).
• Coarse control mode: In this mode, the ADCS sensors idle - the power is turned off for most of the ADCS elements to save energy for the active orbit phases. Six solar cells are being used as SAS (Coarse Sun Sensors), their output is sent to the SCU where the coarse attitude control is taking place.
Figure 2 shows the block diagram of the ADCS. The coarse attitude control system consists of the SCU (Satellite Central Unit), the GAS (Geomagnetic Aspect Sensor) and the 3-axis MTQ (Magnetic Torquers). In this system, the geomagnetic field is measured using the GAS, and the body angular velocity damps using the MTQs.
The fine attitude control system consists of ACU (Attitude Control Unit, 3-axis RW (Reaction Wheel), 3-axis Gyro (Gyroscope), 2 star sensors (HSS), and 4 sun sensors SASH ( Sun Angle Sensor – High precision). In this system, the body angular velocity is measured using Gyro. The quaternion is detected using the HSS. The ACU calculates the control torque and sends the command to the RW. The ACU acquires the current time from the SCU, and sends the internal data as status data and telemetry to the SCU. The commands sent from the ground station via SCU are processed using the ACU for the changing of the target direction or the attitude control parameter.
Figure 2: Block diagram of the ADCS (image credit: Rising-2 partners)
EPS (Electrical Power Subsystem): EPS uses surface-mounted GaAs solar cells for power generation (efficiency of 24%). Four solar panels are mounted on the circumference of the spacecraft (each 8 series x 4 parallels), and one solar panel is mounted on the top side (8 series x 2 parallels). An average power of 42 W is provided. The 9-cell NiMH battery has a capacity of 3.7 Ah (10.8 V). During observations, the spacecraft requires 28.4 W of power while only 7.8 W are needed during non-observation periods. 14)
From the lessons learned of SpriteSat, the EPS has been newly designed and tested. The new design uses an idea of peak power tracking with careful understanding of the characteristics of the solar cells and NiMH batteries.
C&DH (Command and Data Handling) subsystem: The 4 controller units in the bus system are the SCU (Satellite Central Unit, ACU (Attitude Control Unit), SHU (Science Handling Unit), and the PCU (Power Control Unit). The SCU is the main unit of the C&DH subsystem. The ACU, SHU, and the PCU are connected to SCU like spokes. Each unit features an FPGA (Field Programmable Gate Array) and CPU (Figure 3). The SCU has two FPGAs (antifuse and flash) and a CPU. The peripheral units are a magnetometer (GAS), sun sensors (SAS), a GPS receiver (GPS-R), a MEMS attitude measurement unit (TAMU), a U-band receiver (URX), and a S-band transmitter (STX). The antifuse FPGA in SCU has the only reboot function of flash FPGA.
The ACU has a flash FPGA and a CPU. The peripheral units are star CCD sensors (HSS-1,2), a gyro sensor (GYRO), high-precision sun sensors (SASH-1,2,3,4) and a reaction wheel unit (RW). The image taken by HSS is processed in the embedded CPU included in the FPGA of the ACU, and the attitude can be determined. The attitude sensors except HSS are connected by a CAN bus interface. 15)
Figure 3: Block diagram of the C&DH subsystem (image credit: Rising-2 partners)
The SHU has a flash FPGA and a CPU. The peripheral units are the telescope CCD sensors (HPC-R,G,B,M) and the sprite CMOS sensors (LSI-W,N), a fish-eye CCD sensor (WFC), a bolometer array sensor (BOL), and a VLF receiver (VLF-R).
The PCU has an antifuse FPGA without a CPU. The functions are battery charge / discharge control, and peak power tracking (PPT) control for power generation of the solar cells. The control parameters can be modified by uplink commands.
All control units feature A/D (Analog-to-Digital) converters, and the house-keeping (HK) data such as voltage, current, and temperature, are measured in each unit. The status data are sent to the SCU.
Table 1: Summary of FPGA and CPU on Rising-2
RF communications: The S-band is used for downlink data transmissions at 38.4 kbit/s (max). The uplink data rate is 1.2 kbit/s in UHF. The Sendai station, located on the campus of the Tohoku University, is being used for the uplink and downlink services. Further downlink stations are located in Kiruna, Sweden and in Thailand.
Figure 4: Schematic views (CAD models) of the internal configuration of Rising-2 (image credit: Rising-2 partners)
Figure 5: Internal view of the central pillar structure (image credit: Rising-2 partners)
Figure 6: Overview of the spacecraft subsystems and their relations among themselves (image credit: Rising-2 partners)
Figure 7: The sensor complement instruments and their connection to the SHU (image credit: Rising-2 partners)
Table 2: Summary of spacecraft parameters
Figure 8: Photo of the Rising-2 research team at SRL (Space Robotics Laboratory)with their microsatellite (image credit: Tohoku University)
Launch: The Rising-2 microsatellite was launched as a secondary payload on May 24, 2014 (03:05 UTC) on a H-IIA F24 vehicle (No 24) from the Yoshinobu Launch Complex at TNSC (Tanegashima Space Center), Japan. The launch provider was MHI (Mitsubishi Heavy Industries, Ltd.). The primary payload on this flight wa the ALOS-2 spacecraft of JAXA. 16)
The secondary missions manifested on the ALOS-2 mission by JAXA were: 17)
• SPROUT (Space Research on Unique Technology), a nanosatellite of ~7 kg of Nihon University, Tokyo, Japan.
• Rising-2, a cooperative microsatellite (43 kg) project of Tohoku University (Sendai) and Hokkaido University, Sapporo, Japan.
• UNIFORM-1 (University International Formation Mission-1), a microsatellite (~50 kg) of Wakayama University, Wakayama, Japan.
• SOCRATES (Space Optical Communications Research Advanced Technology Satellite), a microsatellite (~ 50 kg) mission of NICT (National Institute of Information and Communications Technology), Koganei, Japan.
Figure 9: Photo of the secondary payloads integrated on the adapter ring of the second stage (image credit: JAXA)
Orbit: Sun-synchronous near-circular sub-recurrent orbit, altitude = 628 km, inclination = 97.9º, period = 97.4 minutes, revisit time = 14 days, number of orbits/day = 15 3/14, LSDN (Local Sun time on Descending Node) = 12:00 hours ± 15 min.
• The Rising-2 spacecraft and its payload are operating nominally in December 2014. 18)
Figure 10: Rising-2 image around Osaki City, Miyagi Prefecture, acquired on Nov. 27, 2014 (image credit: Rising-2 partners)
Figure 11: Cpomparison of NDVI scence acquired by Landsat-8 scence with that acquired of Rising-2 (image credit: (image credit: Rising-2 partners)
• On July 2, 2014, the Rising-2 project succeeded in observing a detailed landscape in sunny spells during the rainy season. Figure 12 is an image of size 3.2 km x 2.2 km at a resolution of ~5m, acquired to the south in Minamiuonuma City in southern Niigata Prefecture. 19) 20)
Figure 12: HPT sample image with the HPT acquired on July 2, 2014 (image credit: Rising-2 partners)
• May 31, 2014: Rising-2 has succeeded in daylight imaging of cloud patterns, and night-time imaging, with a fish-eye CCD camera (WFC) as shown in Figure 13. For comparison, the image is shown side-by-side with a visible image taken by the weather satellite MTSAT (Himawari). The same cloud formation is observed in both images. A circle with a radius of 1,000 km is added as a guide to the meteorological image. 21)
Figure 13: WFC image (right) acquired on May 31 during the day at . (image credit: Rising-2 partners).
Sensor complement: (HPT, LSI-1, LSI-2, WFC, VLF antenna & receiver, BOL)
The sensor complement consists of five scientific instruments and some subsystems: two CMOS cameras with different color interference filters, a CCD camera with fish-eye lens, and a VLF radio wave receiver.
The relationship diagram of sensor complement is shown in Figure 15. Five lenses and a mirror are exposed to the outside of the spacecraft. The newly developed items are HPT and BOL. The other instruments have been already developed in SpriteSat and other previous projects.
HPT (High Precision Telescope):
The HPT is being used to collect the incoming radiation for a particular observational target region. In the case of the sunlit orbit phase, the HPT is generally pointed in the nadir direction to collect high-resolution surface imagery. During the eclipse phase of the orbit, the HPT will be pointed toward targets of opportunity. The main objective is to observe TLEs (Transient Luminous Events).
Figure 14: Illustration of the HPT device (image credit: Rising-2 partners)
The HPT instrument specifications are: (Ref. 18)
- Cassegrain reflective telescope with an aperture of 100 mm and a focal length of ~ 1 m (f/10).
- Compact low mass (3.4 kg) telescope with an aperture diameter of 38 cm
- Four CCD detector arrays are employed in the spectral ranges of 400-650 nm (RGB observations) and of 650-1000 nm. The electronic LCTF (Liquid Crystal Tunable Filter) detection technique is being used for the 650-1000 nm spectral range. This LCTF detection concept, which is polarization sensitive, makes it possible to measure the optical properties of solar radiation reflected from land and sea surfaces. The LCTF can be tuned to any desired wavelength by a computer command within its spectral range.
- CCD resolution: 5 m/pixel. The IFOV generates an image size of 3.3 km x 2.5 km from an orbit of 700 km.
- Spatial resolution: 5-50 m in LCTF mode with a 10 nm band width.
- Exposure time: minimum of 1/4000 second in RGB mode
- The 4 CCD detectors of HPC-R, -G, -B, and -M (Multispectral) were already developed in the SpriteSat project. The detectors have a high sensitivity corresponding to ISO 8000. In addition, they feature an electronic shutter with an exposure time of 1/4000 second.
Figure 15: Overview of the sensor complement (image credit: Rising-2 partners)
The primary and secondary mirrors of HPT feature ZPF (Zero-expansion Pore-Free) ceramics of Nihon providing low-mass and high-strength characteristics. A new grinding technology is being used in the polishing of the mirror surfaces.
Figure 16: Optical design of HPTs for RISING-2 and RISESAT (image credit: Hokkaido University)
The LCTF device is modified for use on a spaceborne imager. The device is comprised of liquid crystal multi-layer plates. The wavelength is variable with narrow 5 nm bands. The central wavelength can be tuned in the range of > 300 nm with 10 ms.
Figure 17: Photo of the world's first spaceborne LCTF (image credit: Hokkaido University)
Why introduction of LCTF (Liquid Crystal Tunable Filter) technology ?
• Reduction in size, mass and power consumption makes it possible to install on a wide variety of platforms for observation
• A multispectral LCTF sensor is comparable to a heavy, expensive hyper-spectral sensor
Figure 18: Schematic view of conventional filter wheel and LCTF technology (image credit: Hokkaido University)
LSI-1 (Lightning Spectrum Imager-1):
The objective is to detect lighting flashes. The camera features a CMOS detector with a format of 512 x 512 pixels. Observations are being made in the spectral band of 744-826 nm. LSI-1 and LSI-2 have a square FOV of 29º corresponding to a ground surface side length of 342 km (Figure 20).
LSI-2 (Lightning Spectrum Imager-2):
The objective is to detect sprites. The camera features a CMOS detector with a format of 512 x 512 pixels.. Observations are made in the spectral band at 762 nm.
Figure 19: Instrument photos of the LSI-1, -2 (left) and WFC (right), image credit: Tohoku University)
Figure 20: Observation scenario of SpriteSat/Rising-2 (image credit: Tohoku University)
WFC (Wide Field-of-view Camera):
The camera is being used to determine the location of lightning flash which is relating to the TGF (Terrestrial Gamma-ray Flashes) event. This high sensitivity panchromatic CCD camera with a fish-eye lens covers a FOV of 140º.
VLFR (Very Low Frequency Receiver):
The objective of (VLF-ANT, VLFR) is to detect TLEs (Transient Luminous Events).
BOL (Bolometer array camera):
The BOL detector array offers observations in the spectral region of 8-14 m, corresponding to the MWIR (Midwave Infrared) and the TIR (Thermal Infrared) regions. At the 700 km altitude, the spatial resolution is ~ 1 km, which corresponds to 0.076º/pixel. Temperature distribution imagery of such targets as: a cumulonimbus region, a ground surface scene, and a region of the sea surface can be generated.
The following three phenomena can be recognized:
- From the temperature of the top of cumulonimbus, the altitude can be estimated. With the simultaneous observations by LSI and WFC, the relationship of transient luminous events and cumulonimbus is analyzed.
- Observing the temperature distribution of the ground surface, buildings, and the sea, the generation of cumulonimbus can be monitored, which is a resource for determining heavy rain.
- Natural disasters such as wildfires and volcano eruptions will be observed; the objective is to check the applicability of this observation method for rapid detection results.
The commercial bolometer camera is being slightly modified for space observations. The instrument has a power consumption of 8.4 W and a mass of 0.554 kg. During a BOL observation mode, the satellite is rotated in such a way as to avoid solar radiation entering the BOL optics. In the rotation period of the satellite, the power of BOL is turned on, and deep space imagery is taken for calibration purposes. After the attitude has been stabilized and the instruments are pointing into target region on Earth's surface, observations may be started.
Figure 21: Photo of the MWIR/TIR bolometer array (image credit: Tohoku University)
Figure 22: Photo of the commercial BOL camera (image credit: Tohoku University)
The observation modes of the spacecraft are being conducted in 15 minute periods, where one period is in the sunlit phase of the orbit and the second observation period is in the eclipse phase of the orbit. One of the following 7 modes is being selected and completed in < 15 minutes. In the EPS, the sensor power is defined as 3 W on average. This requires some adjustment of the time for Mode-1 and Mode-5 to avoid a battery degradation.
1) Mode-1: Sprite observation mode (LSI-1,-2, and VLFR), 3.8 W, only in the eclipse phase
2) Mode-2: Lightning observation mode (WFC, and VLFR), 2.8 W, only in the eclipse phase
3) Mode 3: LSI mode (LSI-2), 1.0 W, for cumulonimbus, ground, and sea, generally in the sunlit phase
4) Mode 4: WFC mode (WFC), 1.0 W, for aurora, ground, and sea
5) Mode 5: BOL mode (BOL), 7.8 W, for cumulonimbus, ground, and sea
6) Mode-6: RGB telescope mode (HPC-R, -G, -B), 3.0 W, for cumulonimbus, ground, moon, and planets
7) Multispectral telescope mode (HPC-M, LCTF), 2.0 W, for cumulonimbus, ground, moon, and planets.
The Tohoku University ground station will be used as the primary station for spacecraft operations. In addition, the ground stations in Kiruna (Sweden) and in Thailand will be used as receiving stations only. The three stations will provide a daily contact time with the spacecraft for ~ 165 minutes.
Figure 23: Overview of the ground station network for Rising-2 operations support (image credit: Tohoku University)
MEVIµS (Model-based Environment for Verification and Integration of µ-Satellite)
SRL (Space Robotics Laboratory) of Tohoku University developed MEVIµS in parallel with the development of RISING-2 and RISESAT. In this environment, all satellite subsystems are simulated in software based on actual components including attitude control as well as data handling and power control. MEVIµS is based on a realtime OS to realize realtime simulation of HILS (Hardware-in-the-Loop Simulation) environment. HILS of the RISING-2 attitude control system was demonstrated by utilizing the satellite on-board computer and reaction wheels. The star tracker test system was introduced to include the failure detection and time delay. As the results, the system development became more efficient since the attitude control system verification could be carried out all of the time. 22)
From the acquisition of the RISING-2 data in on-orbit operations, the reliability of the development environment will be increased by reflecting the results. Furthermore, it will contribute to produce next micro satellite under development. This paper puts an emphasis on the configuration and capabilities of the HILS environment including star tracker test system.
MEVIµS is composed to seven parts: Operation Monitoring System, SSES (Satellite and Space Environment Simulator), EGSE (Electrical Ground Support Equipment), MFE (Modular Front End), Power Supply System, and DCLS (Dynamic Closed Loop Simulator). Figure 7 shows the general view of MEVIµS. All system except the EGSE and the satellite system are connected by a local Ethernet for high speed communication. The SSES and MFE are described which are core elements of MEVIµS.
Outline of SSES: SSES is a software simulator in C++ which constitutes the core of MEVIµS. SSES is composed of a space environment model and a satellite system model, which are separated to maintain the versatility of the simulator.
Figure 24: Schematic of the SSES simulation flow (image credit: Tohoku University)
These models are formed by multiple C++ classes. The former model consists of an orbital calculation class,an attitude calculation class, a geomagnetic model class, a solar model class, and a lunar gravity model class. The latter model includes a satellite mathematical component class, for example, an on-board computer class, a sensor class and so on. Each component class is built in a hierarchic structure to match the actual components. This model is exchangeable to adapt for various satellite systems. If another satellite system model is prepared, all simulation can be implemented to utilize the same space environment model. The calculation flow of SSES and the data transfer model are shown in Figure 24.
The operating frequency of SSES is adjustable. The fastest frequency is 20 Hz from an operational test result. SSES is a mathematical simulator including time delay functions of each component (Figure 25). For an ideal mathematical simulation and time delay simulation, the SILS (Software In-the-Loop Simulation) environment consists of the SSES and Operation Monitoring System. In order to establish HILS, the MFE (Modular Front End) is needed.
Figure 25: Time delay and data transfer model of SSES (image credit: Tohoku University)
Outline of MFE (Modular Front End): MFE a data communication system between the satellite hardware components and the SSES. This system is necessary to execute HILS. The MFE is mainly composed of a controller and a data communication interface. To maintain the versatility, because specific IDs for each satellite are prepared, HILS can be constructed without changing the communication protocol of the system. However, electrical interfaces are need ed for each satellite system. The controller of the MFE features a computer-installed realtime OS, to realize the time management in the microsecond range.
Figure 26: Connection between SSES and MFE (image credit: Tohoku University)
Star Simulator System: The star simulator was developed to test the performance of star trackers in laboratory on the ground. The star simulator overview is shown in Figure 27. Star maps are displayed on an LCD monitor, while the star tracker detects the stars and calculates attitude referred star catalog. The star catalog for the displayed stars can be selected from the Smithsonian Astrophysical Observatory Star Catalog, the Hipparcos catalog, et.al. The star positions on the celestial sphere are calculated from right ascension, declination, and rotation angle of the detector.
Figure 27: Illustration of the star simulator elements (image credit: Tohoku University)
Figure 28: Block diagram of the MEVIµS ( (image credit: Tohoku University)
Attitude control strategy of RISING-2: Attitude determination systems are divided on fine and coarse attitude determination. The sun sensor and magnetometer data is used for the coarse attitude determination. The star tracker data is used for the fine attitude determination. These systems can be changed automatically if roll and pitch angle error fall below 5º and the yaw angle error is < 10º. As might be expected, it is possible to change these systems by using the change command.
The attitude control modes of RISING-2 are the following:
- NP (Nadir Pointing) mode
- TP (Target of earth-surface Pointing) mode
- DSP (Deep-Space Pointing) mode
At the beginning of pointing control, RISING-2 executes the NP control by the coarse attitude determination in order to stabilize the attitude and enable to detect quaternion by star tracker. RISING-2 keeps a free-rotation state for the non-observation time, because the power capacity of RISING-2 is limited. After stabilizing in NP, the accuracy attitude determination system is selected for improvement of the pointing precision, and the attitude control mode is changed as to observation purposes. In TP mode, the target position is sent from the ground station.
In DSP mode, the observation plane of the satellite is pointed to the light avoidance direction without sunlight, moonlight, and the light reflected from the Earth. This mode is carried out for infrared observation using BOL (Bolometer array). After the DSP mode, NP or TP mode, control is executed while sunlight and moonlight do not enter BOL.
In summary, a star simulator was introduced to MEVIµS for the improvement of the HILS environment. To establish this evaluation environment, it is possible to evaluate attitude control system including the actual characteristic of the star tracker. Moreover, this simulator Verification of the RISING-2 attitude control subsystem was demonstrated by utilizing a pre-flight model in the HILS environment prior to the launch. The attitude control procedures were confirmed and the results contributed for making the actual operation method.
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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 (email@example.com).