SDS-1 (Small Demonstration Satellite-1)
SDS is a JAXA (Japan Aerospace Exploration Agency) program initiated in the spring of 2006. The overall objective is to demonstrate on small platforms new technologies in JAXA's Space-technology Demonstration Research Center (SDRC). The program demonstrates advanced space technologies in orbit, at Technology Readiness Level (TRL) 3 to 5, and increases those TRL to space-operation levels, using quick, lowcost systems. 1) 2) 3) 4)
The various SDS missions are selected for parts/components level, system/subsystem level, and mission-concept level, based on importance and urgency in JAXA's satellite technology roadmap. The development of the SDS-series satellites will be a collaborative effort of small and medium sized space companies and JAXA, keeping essential design authority and technologies to be selected for demonstrations by JAXA.
Figure 1: Overview of the SDS series program roadmap (image credit: JAXA)
Figure 2: Artist's view of the deployed SDS-1 spacecraft (image credit: JAXA)
The SDS series spacecraft are designed and developed at JAXA/IAT (Institute of Aerospace Technology). The satellite body is box-shaped, with two solar panels to be deployed in orbit. The requirements call for adaptability to various missions and components. In addition, timely demonstration requires a very short delivery time with low development cost. Overall, the SDS-1 concept builds on the heritage of the MicroLabSat (Dec. 14, 2002 - Sept. 27, 2006) bus concept and technologies by dividing the satellite into bus bay and mission bay. The SDS spacecraft series provides a mass range from 50 -200 kg to enable frequent launch opportunities as a piggyback payload. 5) 6)
Table 1: Main characteristics of the SDS-1 mission
Table 2: Key functions/services required of the SDS bus
Figure 3: Block diagram of the SDS-1 spacecraft (image credit: JAXA)
Figure 4: Mission equipment of SDS-1 (image credit: JAXA)
Launch: A launch of SDS-1 took place on January 23, 2009 as a secondary payload to GOSAT (primary) of JAXA (H-IIA launch vehicle from the Tanegashima Space Center, Japan). The seven secondary payloads on this flight are: 7) 8)
- SDS-1 (Small Demonstration Satellite-1) of JAXA (~100 kg)
- SOHLA-1 (Space Oriented Higashiosaka Leading Association-1), Japan (50 kg)
- SpriteSat (Tohoku University), Japan (microsatellite of ~50 kg)
- PRISM (Picosatellite for Remote¿sensing and Innovative Space Missions) of ISSL of the University of Tokyo, 5 kg
- Kagakaki (SORUNSat-1), Japan, 20 kg
- KKS-1 (Kouku Kosen Satellite-1) of Tokyo Metropolitan College of Industrial Engineering), nanosatellite of 3 kg
- STARS-1 (Space Tethered autonomous Robotic Satellite-1) of Kagawa University, Japan, ~ 10 kg.
Figure 5: Schematic view of the secondary payloads (image credit: JAXA)
Orbit: Sun-synchronous circular orbit, altitude = 666 km, inclination = 98º, LTAN (Local Time at Ascending Node) at 13:00 hours.
• The SDS-1 operations were terminated on Sept. 8, 2010, ending a most successful operational period of about 18 months. All objectives of the mission were demonstrated. 9)
• The spacecraft and its payload are working nominally in 2010. Many lessons were learned in operating the spacecraft, such as: (Ref. 4)
- RF interference with main satellite in critical phase: In critical phases, the project was not able to operate the satellite in several passes due to the RF interference with the main satellite. Frequencies of satellites were isolated, but they were within the AGC (Automatic Gain Control) frequency band of the ground stations. In addition, the output power of the main satellite was +10 dB in the critical phase so that the amplifier of the receiver was saturated and the AGC gain of the SDS-1 signal was depressed.
Figure 6: RF interference with main satellite (image credit: JAXA)
- Advanced operations not planned in advance: After mission success, there are many advanced operation proposals to utilize the satellite. Some of them require processing ability which exceeds on-board computer’s capacity. The project has to define and test maximum processing ability in ground.
- Communication link disconnection in spin attitude: Communication link was sometimes disconnected due to the NULL point of antenna pattern. It is difficult to avoid this in spin stabilized satellite, but we have to conduct analysis considering spin configuration.
• In mission operations (demonstrations), all components are working well, and obtaining excellent results. All mission objectives achieved full success in their nominal operation phase, and SDS-1 operation moved to an upgraded utilization phase which started September 8, 2009. 10)
Table 3: Mission operation results in the nominal operation phase
The following advanced technologies/experiments are to be demonstrated on SDS-1:
1) MTP (Multi-mode integrated Transponder)
2) SWIM (SpaceWire Interface demonstration Module)
3) AMI (Advanced Micro processing In-orbit experiment equipment)
4) TFC (Thin Film Solar Cell)
5) DOS (Small Dosimeter)
6) Small satellite bus technology experiment.
MTP (Multi-mode integrated Transponder):
The next generation MTP improves the operational availability in the satellite TT&C (Tracking, Telemetry and Command) subsystem. The requirements call for a better flexibility of satellite operations than is available in conventional systems. The transponder profile also requires the following improvements: 11) 12) 13)
- Higher information rates
- Small size, low mass, and low power consumption
- Multi-mode capability
- Low cost.
To satisfy these requirements, JAXA is developing MTP in cooperation with NEC. It will be a next-generation S-band standard for various kinds of JAXA LEO/GEO satellites and other operations or missions. MTP provides four modulation schemes (uplink, downlink modulation) as shown in the following list.
1) Use of (PSK/PM, PSK/PM) scheme for the critical phase of satellites after rocket separation
2) Use of the (UQPSK, SQPN) scheme for intersatellite communications
3) Use of the (QPSK, QPSK) scheme for higher information rates
4) Use of the (UQPSK, UQPSK) scheme for simultaneous operations among several satellites
In addition, the new MTP provides the following features:
- Compliance with the CCSDS (Consultative Committee for Space Data Systems) standards and the SFCG (Space Frequency Coordination Group) recommendations. All applicable CCSDS and SFCG recommendations are taken into account.
- Flexibility: MTP features great flexibility in terms of its available modulation schemes, bit rates and ranging methods. It provides several configuration options such as different power levels and interface compatibility with different platforms.
- Introduction of receiving signal discrimination and of an automatic modulation scheme shift function.
- Coherent / incoherent switching: MTP provides both coherent and incoherent performances in four modulation schemes and prohibits an automatic coherent performance, which makes satellite operations flexible.
Figure 7 shows the block diagram of MTP. For a better understanding: in the (UQPSK, SQPN) scheme the uplink signal is called “UQPSK1”; while in the (UQPSK, UQPSK) scheme, the uplink signal referred to as “UQPSK2”.
• The MTO implementation adopts heterodyne circuits in the receiver section. In the heterodyne circuits, a receiving signal is converted into the intermediate frequency (IF) band using a local signal generated in this section. Subsequently, it is converted into a digital signal through an A/D converter and demodulated.
• In the transmitter section, the IF signal is converted into the 2.2 GHz band using a local signal with amplification. The power levels are set by changing the variable attenuator level in a power amplifier component. In the (PSK/PM, PSK/PM) scheme, two power levels can be selected by command.
• The digital signal processing section features carrier synchronization, pseudo-random noise (PN) code synchronization, command data demodulation, telemetry data modulation, and operation mode change functions. All functions are realized in a large size gate array (G/A) chip.
• The electric power section uses a power amplifier component power supply system and other components’ power supply system. Each system comprises an input filter, a current limiter, and a DC/DC converter. It converts an unstable 50 V voltage level into stable voltage levels and supplies some circuits and devices.
The MTP device has a nominal mass of 3.3 kg, a size of 284 mm x 192 mm x 110 mm, a nominal power consumption of < 31 W, and an operational temperature range of -20º to 55ºC.
Figure 7: Block diagram of MTP (image credit: JAXA)
Figure 8: Illustration of the MTP device (image credit: JAXA)
Figure 9: SDS-1 transmission link scenario (image credit: JAXA)
All functions of the MTP device worked very well throughout the mission. This transponder will be used in JAXA’s future King-size satellites (Ref. 9).
SWIM (SpaceWire Interface test Module):
In the timeframe 2005/6 JAXA decided to adopt SpaceWire as a standard reference architecture to be implemented in its future scientific satellites. SWIM uses JAXA's high speed microprocessor as its core CPU, also referred to as SpaceCube 2. The objective is to obtain a scalable mission data handling system and to support the following experiments: 14)
• To demonstrate the concept of the next generation data handling subsystem based on the SpaceWire standard.
• To test the SpaceWire protocol on a chip
• To demonstrate the T-kernel real-time OS (Operating System) and to test standard middleware applications with the T-engine.
• To provide all needed service functions to a super high memory access control using a SpaceWire serial interface link.
Space Cube 2 is a generic multi-mission platform developed by JAXA/ISAS and NTSpace in 2005 for space application. Space Cube 2 is based on the concept of Space Cube 1 and integrates in a single modular stack. Space Cube 2 consists of three modules, the CPU, a data recorder and power supply modules. HR5000, which has been developed by JAXA, is the central processor chip of the system. The chip contains 64 bit microcontroller based on the MIPS 5kf architecture with maximum clock speed of 200 MHz With integrated peripheral devices the chip enables high-speed communication and control. 15) 16)
Figure 10: Illustration of the SWIM breadboard model (image credit: JAXA)
Table 4: Specifications of Space Cube 2
Figure 11: Architecture of the SpaceCube (image credit: JAXA)
Figure 12: Block diagram of SpaceCube 2 (image credit: JAXA)
Figure 13: Photo of the SpaceCube 2 data handling system (image credit: JAXA)
AMI (Advanced Micro processing In-orbit experiment equipment):
The objective is to test the new DC-DC converter (the developed FET is inside the converter) and microprocessor in a space-radiation environment - to prove that the quality and performance of these new parts lives up to expectations. The microprocessor operates at 200 MHz (320 MIPS) to provide on-orbit, high-speed operations support for future applications. 17)
Figure 14: Microprocessor (left) and DC-DC converter (image credit: JAXA)
Figure 15: Block diagram of AMI (image credit: JAXA)
Figure 16: Photo of the AMI device (image credit: JAXA)
TFC (Thin Film Solar Cell):
The objective is to test and verify the mounting technology of two TFC types to be used for next-generation missions. These are:
• Two-junction high efficiency thin-film solar cells
• Flexible Cu (InGa) Se2 (CIGS) solar cells
The in-orbit data will be compared with the performance-prediction model derived from ground-test data.
Figure 17: Thin film solar cell experiment (image credit: JAXA)
DOS (Small Dosimeter):
The objective is to measure the space radiation with a RadFET (Radiation sensitive Field Effect Transistor) located near the semiconductor devices. Highly sensitive RadFETs will be developed, and attached to several points in the SDS-1. The orbit data will be compared with a radiation shield model to evaluate the design of highly sensitive RadFETs. At the same time, the tolerance level of each component to total dose can be verified. 18) 19) 20)
DOS (of SOHLA-1 heritage) consists of RCC (RadFET Control Circuit) and seven sensors. One sensor contains an 8-pin LCC RadFET and a temperature sensor. The RCC consists of two printed boards an aluminum chassis of size: 146 mm x 25 mm x 146 mm. The mass is < 0.6 kg. The power consumption of the RCC is < 2 W. The RCC requires levels of +5 V. The telemetry data interfaces with CCU (Central Control Unit) using the serial RS-422 protocol.
Figure 18: System diagram of DOS (image credit: JAXA)
A RADFET is a specially designed positive channel metal oxide semiconductor (PMOS) transistor with a thick gate oxide; it is optimized for increased radiation sensitivity. The RadFET is suitable for space use in terms of cost, weight, and low power consumption. The RadFET used in the SDS-1 is a 400 nm implanted-gate oxide device manufactured by the Tyndall National Institute in Ireland. For SDS-1, the RadFET bare-die is integrated to an 8- pin LCC package and arranged with a temperature sensor on the printed board. This integrated sensor is 8 mm x 3 mm x 19 mm in size with a mass of ~ 4 g with a 500 mm wire harness (Figure 19).
Figure 19: Photo od a RadFET device (image credit: JAXA)
Figure 20: SDS-1 structure and mission and bus components with regard to DOS (image credit JAXA)
Small satellite bus technology experiment:
The objective is to establish functional readiness for the future mission requirements in the SDS program. Three high-priority components were selected within the development strategy of the next-generation small satellite bus technologies to be demonstrated:
• GPSR (Small GPS Receiver)
• MSS (Micro fine Sun Sensor)
• ACMR (Advanced Monitor Camera)
The GPSR was developed based on car navigation systems. The goal is to meet space requirements. The body structure, GPS antenna and firmware were modified for endurance in space. Radiation testing of some commercial-grade electrical parts confirmed their good tolerance.
The MSS selected offers a good balance with the small satellite's highly functional attitude control system in dimension, mass, power consumption, and performance. Its bias error is less than 0.1º (3σ) and random error is less than 0.01º (3σ). The detector on the MSS is a commercial-grade CMOS / APS (Active Pixel Sensor) device.
The ACMR is a next generation optical Earth imager that uses a state-of-the-art COTS CMOS device. The data handling board will be a next generation small satellite main OBC and will be evaluated in orbit.
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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.