Minimize HORYU-4

HORYU-4 (High Voltage Technology Demonstration Satellite-4) — also known as AEGIS (Arc Event Generator and Investigation Satellite)

Overview   Spacecraft   Launch    Mission Status   Sensor Complement   References

HORYU-4 is a nanosatellite mission of Kyushu Institute of Technology (Kyutech), Fukuoka, Japan. Kyutech is located in the Kitakyushu region, about 1000 km southwest of Tokyo. The objective of HORYU-4 is to acquire on orbit data of discharge phenomena occurring on high voltage solar array to deepen understanding of satellite charging, to contribute to the reliability improvement of current space systems, and to positively contribute to the realization of future high power space systems. 1)

HORYU-4's main mission is to acquire an arc current waveform by an onboard oscilloscope and capture its image by a camera triggered by the oscilloscope. In addition, HORYU-4 plans to carry out scientific experiments on arc-mitigation high voltage solar array, plasma measurement using a double Langmuir probe, vacuum arc thruster, photo-electron current measurement, and polymer material degradation.

ESD (Electro-Static Discharge), the arcing phenomenon occurs mainly on solar array panels and can cause partial or even complete satellite power loss, which often results in the inability to perform the missions and ultimately in the satellite loss. From 1997 to the middle of 2000s, ESD on power system, such as solar array, solar array paddle drive motor, and power harness, was the major failure cause of satellites, especially for high power GEO telecommunication satellites that adopted 100 V bus to answer the demand of higher power usage due to proliferation of satellite broadcasting. 2)

As a result of international collaborative efforts to improve the reliability of satellite solar panel and power system, ISO-11221, "Space systems — Space solar panels — Spacecraft charging induced electrostatic discharge test methods" was published on August 1, 2011. Although the testing methods were developed with the best knowledge of what happens in space, there is nobody who saw where a discharge occurs on a solar panel nor its discharge waveform. No tests has ever been carried out in space to make sure that what is observed in laboratory is actually the same as what is happening in space.

The first motivation of HORYU-4 project is the desire to answer questions such as: What are the effects of the difference between the plasma and vacuum parameters in chamber vs. space? Do chamber walls influence results? Has satellite motion a critical role in the observed results? These questions need to be answered before confirming or disconfirming the theory and ground-based observations the standard is based on. The results obtained by the HORYU-4 project will be reflected to the revision process of ISO-11221. 3)

The second motivation is the aspiration to actively contribute to the development of the next generation high power space systems by providing a safe and reliable way of generating high photovoltaic power. In the 2020's, the International Space Station plans on retiring, and future projects such as planetary exploration, space solar power systems, and space hotels are considered. These systems require high power, 1 MW in the near future, and if their bus voltage stays at a safe level of about 100 V, their weight will be tremendously high, whereas if the bus voltage could be safely raised to 300 V, the future high power space systems mass could be reduced of nearly 20%. 4)

The HORYU-4 project started in May 2013 after securing funding from JSPS (Japan Society for Promotion of Science). It was selected as a H-IIA secondary payload with Astro-H ("Hitomi") in July 2014.

Regarding the project structure, the HORYU-4 project took a three-stage structure, unlike HORYU-2 where only two stages existed, i.e. professors and students. To assure the success of the minimum mission, professional staff was hired. Then the team structure was professors-staff, and students. The project manager was of the professional staff. This hierarchy sometimes posed confusion and difficulty in the chain of commands.

This difficulty was partly produced by the large number of team members, more than 30. There are two reasons why we had so many team members. The first is the satellite carried so many mission payloads. The second is that we tried to utilize the HORYU-4 development for education of students from non-spacefaring countries. The large team size caused problems in intra-team communication, which was further inflated by the diverse cultural background of the team members.

The lessons learned from these points are that we should simplify the satellite missions and the team structure. There is a Japanese saying that "those who chase two rabbits catch neither". If the objective is research, we should focus on the research and try to maximize its outcome as much as possible without thinking to use the project for education. The number of mission payloads onboard a satellite should be kept to a minimum to avoid complexity in satellite development.

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Figure 1: Photo of the HORYU-4 development team( 46 staffs and students from 18 countries - associated with the Space Engineering International Course,Postgraduate program at Kyutech), image credit: Kyutech 5)

 


 

Spacecraft:

HORYU-4 is a nanosatellite within an envelop of 450 mm x 420 mm x 430 mm with a mass of about 10 kg. The spacecraft bus is based on the HORYU-2 bus and lessons learned from the HORYU-2 failures were implemented especially with regard to single event latch-up instances. Figure1 shows the satellite and its elements as it was attached to the launch vehicle (H-IIA). The UHF antennas, the VHF antenna, and the arm holding the mirror are fixed, which makes HORYU-IV a deployment-free nanosatellite.

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Figure 2: Illustration of the HORYU-4 nanosatellite (image credit: Kyutech)

1. L-band antenna
2. GPS antenna (AOCS)
3. UHF antenna
4. Langmuir probe (DLP)
5. Bus solar array (EPS)
6. High voltage solar array (HVSA)
7. Sun sensor (AOCS)
8. Earth observation camera (CAM)
9. Vacuum arc thruster (VAT)
10. Discharge experiment solar array (DEG)
11. Mirror (AVC)
12. Secret ink (INK)
13. Photo-electron current measurement (PEC)
14. Electron collector (HVSA)
15. VHF antenna
16. UHF antenna
17. S-band antenna

Figure 2 shows how the payload and bus are accommodated.

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Figure 3: HORYU-4 external configuration (image credit: Kyutech)

1. S-band antenna
2. UHF antenna
3. Earth observation camera (CAM)
4. L-band antenna
5. High voltage solar array (HVSA)
6. Langmuir probe (DLP)
7. External connector
8. Sun sensor (AOCS)
9. Bus solar array (EPS)
10. Arc vision camera (AVC)
11. Secret ink (INK)
12. Discharge experiment solar array (DEG)
13. GPS antenna (AOCS)
14. UHF antenna
15. VHF antenna

The three main differences between HORYU-4 and HORYU-2 bus can be found in the OBC (On-Board Computer), AOCS ( Attitude and Orbit Control Subsystem) and COM (communication subsystem). The OBC features three microcontrollers: two H8 that watch each other and one PIC (Peripheral Interface Controller). A power reset of the H8 occurs every 24 hours independently of an anomaly occurrence. In case of H8 malfunction, the PIC can reset the satellite through commands sent from the ground station.

For COM, in addition to UHF/VHF antennas, HORYU-4 also uses an S-band patch antenna for data transmission, and an L-band patch antenna for data reception. Mission data and offline housekeeping data are sent at a frequency of 2400.3 MHz (S-band amateur frequency spectrum) at a rate of 100 kbit/s. In case of an S-band malfunction, the mission and housekeeping data will be sent via the UHF link at a frequency of 437.375 MHz and a rate of 1.2 kbit/s. Beacon and real-time housekeeping data, providing status information are sent via the UHF antenna at a frequency of 437.375 MHz and a speed of 20 wpm and 1.2 kbit/s, respectively. Commands are uplinked via the VHF antenna at a frequency of 145-146 MHz and a rate 1.2 kbit/s. The L-band is used to send a reset signal to the PIC in case of an OBC malfunction using DTMF (Dual-Tone Multi-Frequency).

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Figure 4: Block diagram of OBC-COM (image credit: Kyutech)

HORYU-2 suffered a series of malfunctions due to SEL (Single Event Latchup) of its H8 microprocessors. Fortunately, if the two processors suffer SEL, the increased consumption current depletes the battery and the satellite shuts down itself and recovers afterward. For HORYU-4, various mitigations against SEL were implemented. First-of-all, the threshold value of the overcurrent protection circuit was determined based on the SEL test of the H8 processors, where the current consumption of the increased during SEL was measured. Secondly, a PIC microprocessor (PIC16F876A), that worked continuously onboard HORYU-2, was added to do a power-reset independently if the H8 suffers an SEL. Thirdly, an L-band communication line was added to send a reset signal to the PIC from the ground.

As with HORYU-2, HORYU-4 uses a permanent magnet and hysteresis damper for passive attitude and orbit control. The novelty for HORYU-4 is in the attitude and orbit determination, for which it uses six sun sensors and a GPS receiver in addition to two three-axis control gyro-sensors.

Parameter

HORYU-2

HORYU-4

Spacecraft size (X x Y x Z)

350 x 310 x 315 mm

450 x 420 x 430 mm

Spacecraft mass

7.1 kg

10 kg

Orbit

Sun-synchronous sub-recurrent orbit, altitude = 670 km

Circular orbit, altitude = 575 km, inclination = 31º

Downlink

- UHF (437.375 MHz) CW beacon (20 wpm)
- UHF (437.375 MHz) FM 1200 bit/s for data

- UHF (437.375 MHz), CW, beacon (20 wpm)
- UHF (437.375 MHz), FM, 1200 bit/s for data
- S-band (2400.3 MHz), 100 kbit/s for data

Uplink

VHF (145-146 MHz)

- VHF (145-146 MHz)
- L-band (1.26 GHz)

Solar array power

- Max: 4.5 W
- Average: 2.6 W (measured in orbit)

- Max: 9 W
- Average: 5.2 W

Battery

- Ni-MH
- 3 parallel, 3 series

- Ni-MH
- 3 parallel, 6 series (5700 mAh at 7.2 V)

Power consumption

- Nominal: 0.7 W
- Peak: 3.3 W

- Nominal: 5.1 W
- Peak (all functions ON) 15.3 W

Attitude control

Passive control by permanent magnet and hysteresis damper

Same as HORYU-2

Attitude sensors

Gyro sensor

Gyro sensor, Sun sensor

Orbit determination sensor

None

GPS receiver

Deployment of elements on orbit

UHF/VHF antenna

None

Thermal control

Passive + battery heater

Passive + battery heater

OBC

Renesas H8, HD64F36057FZV x 2

- Renesas H8, HD64F36057FZV x 2
- PIC 16F876A for power reset

Table 1: HORYU-2 and HORYU-4 bus characteristics comparison

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Figure 5: Block diagram of the HORYU-4 nanosatellite (image credit: Kyutech) 6)

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Figure 6: Illustration of the HORYU-4 nanosatellite with the components shown (image credit: Kyutech)

 

Launch: The HORYU-4 nanosatellite was launched as a secondary payload on Feb. 17, 2016 (08:45:00 UTC) on an H-IIA -F30 vehicle from the Yoshinobu Launch Complex at at TNSC (Tanegashima Space Center), Japan. The primary payload on this flight was Astro-H (Hitomi) of JAXA.

Orbit: Near-circular orbit, altitude of ~575 km, inclination = 31º, period of 96.2 minutes.

Secondary payloads:

• ChubuSat-2, a microsatellite of Nagoya University and of Daido University to observe the radiation from Sun and Earth with a radiation detector.

• ChubuSat-3, a microsatellite of Nagoya University and of Daido University to collect AIS signals from ships.

• HORYU-4, a 10 kg technology demonstration nanosatellite of Kyushu Institute of Technology.

 


 

Mission status:

• As of September 8, 2016, the satellite is in healthy condition. Although HORYU-4 was developed based on HORYU-2, the first satellite launched by Kyutech, and its lessons learned, there are many new lessons learned this time (Ref.1) .

- The satellite carries a VGA (640 x 480) color camera both in ±Y surface. The camera is activated by the uplink command. As there is no active attitude control, the Earth is not captured always. In most of the cases, one camera faces the dark space and the other faces the Earth. Figure 7 shows an image taken over Sudan. The Nile river going through Khartoum is clearly seen in the photograph. At this region, the river width of the Nile is about 2.5 km and the river extends 5 pixels or more. HORYU-4 took more than 300 pictures in the first 6 months.

- Several lessons were learned during the operational phase. The effect of satellite attitude on communication should not be underestimated. Also, the Doppler shift can have a critical effect on the communication. Verification of the satellite-to-ground communication is often done by an end-to-end test. The test results should be looked at with caution if the satellite motion is not taken into account.

- Another lesson leaned from the operational phase is how to keep motivation of students and staff who are engaged in operation. Just after the launch, the motivation is very high. But soon it starts diminishing. To keep the motivation, the mission results should be shared among the team members no matter how small they are.

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Figure 7: Photo over Khartoum, Sudan, acquired by the CAM (image credit: Kyutech, Ref. 1)

• On February 24, 2016, the project carried out the first experiment of the main mission. The solar array in +X panel was biased for 30 minutes. At 550 seconds from the start of the experiment, the first discharge was observed. At 834 seconds, the second discharge was observed. Figure 8 shows the discharge waveform captured by OBO (On-Board Oscilloscope) and Figure 9 shows the discharge image captured by the AVC (Arc Vision Camera). The peak current of 60 A and the pulse width 1.5 µs are within the expected range. The discharge location was at the edge of the solar cell. These data is the world premiere on-orbit observation data of discharge in orbit. With the success of this experiment, the HORYU-4 project achieved the minimum mission success. For more details of this experiment, see 7)

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Figure 8: Discharge waveform observed in orbit (image credit: Kyutech)

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Figure 9: Discharge location of the discharge waveform shown in Figure 8 (image credit: Kyutech)

• Figure 10 shows the long-time trend of the battery temperature. The satellite has two battery packs. A temperature sensor is attached to each battery pack. On Day 1 after the launch, the battery temperature 1 was +25ºC. It showed a gradual increase to +40ºC in one week and eventually reached to +80ºC. The project believes the temperature is not due to malfunctions of the sensors. The anomalous high temperature appeared three times with a period of 20 days, approximately. The initial guess was an impedance mismatch between the two parallel battery packs, the current dropped to the low impedance pack. The battery temperature 2, however, also showed the anomalous high temperature in 120 days. Therefore, the project guess is that the satellite attitude and the beta angle have some effects on the periodic temperature rise. Fortunately, the battery packs 1 and 2 are still healthy.

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Figure 10: Long time trend of battery temperature (image credit: Kyutech)

• Figure 11 shows the long-time thermal trend of the +Y surface. There are occasional jumps in the temperature, which the projects believes are due to malfunctions of either the sensor electronics or data acquisition hardware. Despite the occasional anomalous data, one can see that the satellite external panel stays in the temperature range of -15ºC to +50ºC.

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Figure 11: Long time trend of +Y surface temperature (image credit: Kyutech)

• Figure 12 shows a trend of the satellite rotational speed. After 3 months, the satellite rotation was around the X axis, the direction of the permanent magnet. Only minor libration motion was noticed around the Y and Z axis.

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Figure 12: Long-time trend of satellite rotational speed (image credit: Kyutech)

• Figure 13 shows a video frame taken by the rocket second stage after deployment of HORYU-4. From the video, the project determined that the satellite was rotating around Y axis with 15º/s. The rotation has been gradually damped by the internal hysteresis damper.

 

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Figure 13: HORYU-4 after separation. The video was taken from the H-IIA second stage (image credit: JAXA digital archive)

 


 

Sensor/Experiment complement: (HVSA, DLP, VAT, INK, PEC, CAM, SNG)

HORYU-4 is outfitted with several mission payloads and experiments. The primary objective of the mission is the demonstration of High-Voltage Solar Arrays that operate at 300 V. Two arrays of 24 x 10 mm solar cells are installed on the +Y and -Y faces of the satellite. Because of the higher potential difference between the solar array surface and the surrounding plasma mitigation techniques are required to avoid inadvertent discharges. 8)

 

HVSA (High Voltage Solar Array)

The objective of HVSA is to generate a high voltage to induce a 300 V or higher potential difference between solar array surface (triple junction type) and the surrounding plasma to study discharges and their possible mitigation techniques.

 

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Figure 14: Special spherical solar cells to generate 300 V (image credit: Kyutech)

Figure 15 shows circuit diagram of the on-board high voltage experiment system. This experiment system and the satellite main system are electrically isolated including control and measurement signals. The HVSA controller board distributes high voltage from HVSA (i.e. 350 V) to sample solar arrays. The generated high voltage also distributed to the DLP and VAT for surface cleaning and drive, respectively.

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Figure 15: Block diagram of the HVSA (image credit: Kyutech, Ref. 6)

 

DLP (Double Langmuir Probe)

DLP is a novel plasma measurement system designed and developed by LaSEINE (Laboratory for Spacecraft Environment Interaction Engineering) at Kyutech. This mission applies the principle of Langmuir Probe which operates by inserting one or more electrodes into plasma, with a constant or time-varying electric potential between the various electrodes or between them and the surrounding vessel. For this system, a varying voltage is applied between the two probes LP+ and LP- and a measurement system records this voltage and the current as a result to measure plasma parameters (temperature, density, and plasma potential) at the orbit of the HORYU-4 satellite (575km, 31º). This mission shall contribute to the characterization of discharge on solar panels due to high voltage in space with respect to plasma condition.

To measure the density and temperature of the plasma the satellite flies through, HORYU-4 is outfitted with Double Langmuir Probes (Figure 16). Langmuir probes can determine electron temperature, electron densities and the electric potential of a plasma. Two electrodes are inserted into a plasma environment. The electrodes have a constant or time-varying electric potential between them to allow the determination of physical plasma properties by measuring currents and potentials in this two-electrode system. A bias voltage is applied to the probe and the resulting current that is measured is proportional to plasma charge density.

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Figure 16: FM of the HORYU-4 satellite showing the location of DLP in the +Y and -Y directions (gold plated PCB of 4cm×14cm), image credit: Kyutech)

Figure 17 shows the DLP system which consists of probe biasing circuit, the analog measurement circuit, and the controlling electronics. The probe biasing circuit is controlled by ON and OFF signals from the Big Apple microcontroller. The system has 5 switches that perform 3 main functions: capacitor discharge, DLP mission, and probe cleaning. This system is isolated form the satellite bus through an OpAmp isolator. Five photo MOSFET switches are used to accomplish 3 sets of controlling tasks: 1 switch for capacitor discharge, 2 switches for mission, and 2 switches for probe cleaning. The photo MOSFET switch is a solid state relay which consists of a LED (Light-Emitting Diode) for the input side and MOSFETs for the contact point. It is not only smaller and lighter, but also easier to drive and faster than conventional mechanical relays.

Any of the 3 types of signals from the Big Apple microcontroller (capacitor discharge, mission, and probe cleaning) causes a forward current to flow through the photo MOSFET.

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Figure 17: DLP system overview (image credit: Kyutech)

The DLP electronics are located on an electronic board called the Big Apple board as shown in Figure 18. This board also has 3 other missions: VAT (Vacuum Arc Thruster) mission, PEC (Photo Electrons Current) measurement mission, ELF/SCM (Electron-Emitting Film)/Surface Charging Monitor) mission. These four missions share a microcontroller that controls the operations, how power is distributed, and how data are saved. The Big Apple microcontroller receives command from the OBC (On-Board Computer) to carry out any of the four missions and also to downlink missions data. The power to this board is at a voltage of 5 V supply from the EPS (Electric Power System).

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Figure 18: Block diagram of the Big Apple mission board (image credit: Kyutech)

 

VAT (Vacuum Arc Thruster)

The objective is the on-orbit demonstration of a trigger-less vacuum arc thruster, whose circuit is directly connected to the high voltage solar array. The objective of VAT to use this thruster onboard of micro- and nanosatellites for attitude control, orbital station keeping, or as momentum wheels. 9)

The VAT concept consists in generating a vacuum arc thanks to the load accumulated into the condenser. Then, upon vacuum arc creation, metallic vapor is ejected from the cathode and this reaction is used as the thrust force. For charging the condenser, HVSA is used, which makes VAT a direct drive system (300 V directly apply from HVSA system) that does not require booster circuit. Moreover, the cathode used as propellant is made of CFRP (Carbon Fiber Reinforced Plastic). Thanks to the interaction of the CFRP surface with the surrounding plasma, very small discharges can be generated to trigger vacuum arcs. The structure of the thruster head that includes the cathode, anode, and insulator parts is described on Figure 19.

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Figure 19: Schematic of the VAT system (image credit: Kyutech)

In Figure 20, the green circuit line is the charging line, directly connected to HVSA, and the red line is the discharge circuit with a 10 µF main discharge capacitor connected to the Cathode and Anode of VAT. On the cathode probe, the current is used to measure the discharge parameters on-board with the OBO (On-Board Oscilloscope).

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Figure 20: Electric circuit for the HORYU-4 VAT (image credit: Kyutech)

 

INK (Secret Ink)

Study of surface degradation by atomic oxygen. INK is comprised of several different polymer materials that will be monitored over an extended period of time for an assessment of their degradation caused by atomic oxygen. The Photo-electron Current Measurement includes equipment for the generation of electrical spectra on a metal and isolator when illuminated by the sun. A CMOS camera installed on the satellite can deliver color photos of Earth to be used in space awareness campaigns and education.

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Figure 21: Illustration of the INK layout (image credit: Kyutech)

 

PEC (Photo-Electron Current) measurement

PEC is one of the major sources contributing to spacecraft charging. Hence, the mission objective is to measure PEC as a consequence of spacecraft charging. The PEC mission development aimed at establishing a current amplifier able to measure currents of the order of nA (nano Ampere) from metallic and insulator surfaces using the AM0 (Air Mass zero) spectrum.

This mission is based on the principle of photoelectric effect in which the energy carried in a photon is transmitted to an electron in the surface of a material. The electron absorbs the photon's energy and is displaced from the surfaces of the material, thereby creating a small electric current. A current-voltage amplifier circuit is then used for amplification of this current and data collection using a PC. On-board HORYU-4, three different materials are mounted for photo-electron current measurement: Gold, KaptonR (polyimide), and Black KaptonR. Figure 22 describes the mission details.

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Figure 22: On-orbit measurement of photo-electron current (image credit: Kyutech)

This mission is necessary to estimate the photo-electron emission current influence on spacecraft charging, hence the discharge phenomenon. This will help to predict the electric properties of various materials irradiated by sunlight. Thanks to the collected data, for a given application, the most suitable materials could be determined during a spacecraft design phase.

The PEC system is a current-voltage amplifier circuit that consist of LMC6001 ultra-low input operational amplifier, the MCP6032SN operational amplifier, capacitors, resistors, and others discrete components. This circuit is designed to have three sub-circuits mounted on a single board for the measurement of each sample output of photo-electron current.

This circuit was designed with EAGLE (Easily Applicable Graphical Layout Editor) software and the components were located on the printed board as defined in the circuit diagram and then soldered. Figure 23 shows the schematic diagram of the PEC system.

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Figure 23: PEC measurement system schematic (image credit: Kyutech)

The PEC system is connected to the Big Apple board that serves as an interface to the PEC mission execution. The sequence of operations is as follows:

1) The OBC writes a command to the Big Apple flash memory

2) The Big Apple microcontroller reads the command from the Big Apple flash memory and closes a 5 V switch to provide the adequate power to PEC

3) If the sun sensors measurements, provided by the AOCS , show that the incident sunlight on the +Z panel is within ±15º, then the Big Apple microcontroller closes the 24 V switch to bias the PEC grid

4) Mission starts

5) After the mission ends, the Big Apple microcontroller receives data from PEC, which are then written to Big Apple flash memory

6) The OBC reads the PEC data through Big Apple flash memory prior to data transmission to a ground station.

 

CAM (Camera for Earth photography)

The CAM subsystem is responsible for capturing Earth images. It is part of HORYU-4's project outreach activities to communicate with local and international communities. The objectives are:

- To take images of the Earth

- To take images of the Earth at a specified location

- To store images on its own memory for OBC to access and read. The images will be transmitted by the COM subsystem to any ground station.

- To contribute to HORYU-4's outreach activities through Earth imaging.

The CAM subsystem is sharing its PCB (Printed Circuit Board) with the AOCS. The subsystem consists of the following: a PCB board containing the electronics (Figure 24 left); and two camera modules mounted on the satellite Y panels, and a camera modules' harness (Figure 3 right).

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Figure 24: AOCS/CAM board (image credit: Kyutech)

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Figure 25: Left: C1098 camera module (front and back sides); right: Camera module harness (image credit: Kyutech)

CAM modes of operation: Essentially, there are 3 main modes of operation:

1) Timer Mode: Start at time t specified by the OBC. On the timer signal, it shoots an image or several images and saves to flash memory.

2) Normal Mode: The CAM takes one picture of the Earth at a time using AOCS data. If the image taken is not in space, the data will be saved to flash memory.

3) Target Mode: Reads the GPS data, compares the longitude and latitude with the data received by the OBC. Then it takes the number of images at the specified time interval. The images are saved to memory.

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Figure 26: CAM modes of operation: Timer mode (left), normal mode (center), target mode (right), image credit: Kyutech

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Figure 27: CAM system block diagram (image credit: Kyutech)

The CAM starts operating when switched ON by reading the command saved by OBC to CAM-memory. Depending on the command type and parameters, CAM reads the GPS data via Rx serial line connected to the AOCS. CAM executes its mission and saves the image data to the CAM's 8 MB flash memory. OBC then accesses and reads the information onto CAM's flash memory and transmits it to any ground station via COM. The CAM PIC erases the CAM flash memory (all sectors) every time a new command is received from OBC H8 Main.

 

SNG (Digi-Singer)

The SNG is an amateur payload which uses the vocal synthesizer on board the satellite to send songs from HORYU-4 to the ground using the UHF system for space education purposes. SNG can be used by radio operators to upload songs to the satellite, use a vocal synthesizer on the satellite and then send back the song on the UHF frequency.

 


1) Mengu Cho, Pauline Faure, Atomu Tanaka, and the HORYU-e Project Team, "Development Philosophy and Flight Results of Arc Event Generator and Investigation Satellite HORYU-IV," Proceedings of the 67th IAC (International Astronautical Congress), Guadalajara, Mexico, Sept. 26-30, 2016, paper: IAC-16-6B,12

2) Frost and Sullivan, "Commercial GEO Satellite Bus Reliability Analysis, TU Delft, August 2004, URL: http://www.lr.tudelft.nl/organisatie
/afdelingen-en-leerstoelen/department-space-engineering/space-systems-engineering/space-links
/commercial-communications-satellite-bus-reliability-analysis/

3) Mengu Cho, "Systematic Review of ISO-11221," 14th SCTC (Spacecraft Charging Technology Conference), ESA/ESTEC, Noordwijk, The Netherlands, April 4-8, 2016, URL: http://esaconferencebureau.com/custom/
16a04/14th%20SCTC%20Proceedings/Papers/(3)%20Wednesday%206%20April
/(5)%20Session%201%20-%20Standards/1020_Cho.pdf

4) J. R. Brophy, R.Gershman, N. Strange, D.Landau, R. G. Merrill, T. Kerslake "300-kW solar electric propulsion system configuration for human exploration of near-Earth asteroids," Proceedings of the 47th AIAA (American Institute of Aeronautics and Astronautics), San Diego, California, 2011, URL: http://www.kiss.caltech.edu
/workshops/space-challenge2011/references/propulsion-system
/300-kW%20Solar%20Electric%20Propulsion

5) Mengu Cho, Pauline Faure, "Overview of Arc Event Generator and Investigation Satellite HORYU-4," 14th Spacecraft Charging Technology Conference, ESA/ESTEC, Noordwijk, The Netherlands, April 7, 2016, URL: http://esaconferencebureau.com/custom/16a04/
14th%20SCTC%20Proceedings/Presentations/(4)%20Thursday%207%20April/(9)%20Session%204%20-%
20Solar%20Array%20Plasma%20Interactions/1245_Cho.pdf

6) Tatsuo Shimizu, Mengu Cho, "HORYU-4: Miniaturised Laboratory for In-Orbit High Voltage Technology Demonstration," JAXA, SP-15-02, URL: https://repository.exst.jaxa.jp/dspace/
bitstream/a-is/562167/1/AA1630004030.pdf

7) Tatsuo Shimizu, Hiroshi Fukuda, Kazuhiro Toyoda, Mengu Cho, "Solar Array Electrostatic Discharge Current and Image Captured in Orbit," Journal of Spacecraft and Rockets,Publication date (online): October 20, 2016, doi: 10.2514/1.A33622

8) "HORYU-4 payload," URL: http://kitsat.ele.kyutech.a
c.jp/horyu4WEB/payload.html

9) Kateryna Aheieva, Shingo Fuchikami, Hiroshi Fukuda, Tatsuo Shimizu, Kazuhiro Toyoda, Mengu Cho, "Vacuum Arc thruster development for HORYU-4 satellite," URL: https://repository.exst.jaxa.jp/dspace/bits
tream/a-is/549913/1/SA6000036081.pdf

 


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 (herb.kramer@gmx.net).

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