ISAS/JAXA Small Scientific Satellite Series Program with Flexible and Reusable Bus Design

1. 26th ISTS @Hamamatsu [2008-f-17] ISAS/JAXA Small Scientific Satellite Series Program with Flexible and Reusable Bus Design T. Nakagawa (ISAS/JAXA) S. Sakai (ISAS/JAXA)

2. ISAS/JAXA small scientific satellite series • The Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (ISAS/JAXA) has recently started to develop a series of small scientific satellites, and has a plan to launch three to five minisats (300-400 kg) per five years. • The series aims at ‘cheaper and faster realization of unique space experiments’ as a complementary program of mainstream medium-class scientific satellites. The mission payloads are on the roof of the bus system. • In order to shorten the period of satellite development with low cost, it is considered to be reasonable to employ standard bus architecture, where the bus and payloads are clearly separated in a modular manner. • The modular configuration is attained in exchange of resources like mass. Mission payload Standard bus

3. Working Groups for small scientific satellites First flight (2012)

4. This presentation • The missions for small scientific satellites have various kinds of requirements: • three-axis/spin-stabilized attitude control, • a wide range of power (50 to 300 W), • a variety of orbit (LEO, SSO, HEO, etc), ... -> A traditional rigid standard bus will be unable to satisfy the diversity of mission requirements. • In this presentation, several concepts to enhance flexibility of the standard bus are introduced: • Layers of standardization • Semi-custom-made bus system • Tradeoff between integration and modularization

5. Layers of standardization • Specifications of most conventional standard buses are strictly regulated for a particular mission in the whole system/subsystem of satellites. (e.g., communication satellites on GEO) • Effectiveness of the rigid standardization depends on the number of satellites with the particular mission. • Traditionally, medium-class scientific satellites in ISAS/JAXA were dedicatedly developed in order to satisfy various requirements and pursue resource-optimal design, so that fruitful achievements would be brought. • Optimization of satellite-by-satellite consumes relatively large cost and long development time. -> ‘layers of standardization’ concept scientific satellites with a wide variety of requirements

6. Layers of standardization [Cont’d] The Layer (IV): Configuration • Similar to the conventional rigid standard buses • Possible to omit mechanical/thermal model tests • Applicable to most scientific observation missions by adding selectable or alternative options The Layer (III): Instruments • Advantageous for engineering missions where the satellite configuration itself is hard to be standardized (e.g., the small lunar lander with a large-scale propulsion system) • Valuable concerning economics of bundling order The Layer (II): Interfaces • Essential for higher levels of standardization The Layer (I): Design Methods • Standard method of modeling satellites -> reuse of GSE software, etc SpaceWire as standard data network • Promising technology because of high speed, simplicity, and testability • RMAP (Remote Memory Access Protocol) over SpaceWire is a powerful method, which enables one component to read and write directly from memory and registers of another component over the network.

7. Semi-custom-made bus system • Consideration of the bus specifications using a requirement matrix • Three categories: Core, alternative option, and selectable option • The alternative and selectable options are crucial for the flexible standard bus to realize diverse missions with time- and cost-saving strategies. Semi-custom-made bus system For small scientific satellites Ready-made standard bus full-custom-made medium-class scientific satellites • The idea is similar to Product Line or Product Platform engineering in development of commercial products. • The capability of adding or replacing options is based on recent progress in technology of modular design and Plug and Play. • A tradeoff between the number of options and effectiveness of the standardization should be carefully investigated.

8. Semi-custom-made bus system[Cont’d] Selectable options • solar array drive assembly (SADA) • GPS receiver • X-band transmitter • monopropellant propulsion system • design for EMC (Electro-Magnetic Compatibility) Alternative options • number of solar panels (one, two, or three panels per wing) • capacity of Li-ion batteries (35Ah/50Ah) • accuracy of attitude sensors such as a star tracker and an inertial reference unit (corresponding to requirements) • size of reaction wheels (corresponding to requirements of perturbation immunity) Note: Nominal specifications are underlined.

9. Tradeoff between integration and modularization Cost analysis says that present satellite cost has strong correlation with the number of onboard components. -> Common functions which belong to different components have to be integrated into a single component. The integrated architecture is intrinsically appropriate for small satellites because it will consume less resources such as mass, size, and electric power. What is important is to standardize interfaces between modules. -> The standard interfaces will facilitate change of architecture from integration to modularization, and vice versa. The integrated architecture, however, has latent difficulties in testability and reusability. -> Modularization is attempted in the next step. The modular architecture seems preferable in many cases, if a resource budget including cost allows it. Conventional Integration Modularization

10. Tradeoff between integration and modularization [Cont’d] The functional integration is considerably pursued. System block diagram of the standard bus for small scientific satellites (baseline)

11. Tradeoff between integration and modularization [Cont’d] System block diagram of the standard bus for small scientific satellites (modification) Now studying to modify the architecture to more modular one so as to enhance testability and reusability

12. “Semi-order-made” approach[e.g. AOCS] • So, now “semi-order-made” approach is considered. • Example: what is the appropriate attitude & orbit control system (AOCS) configuration to implement this concept. • IRU • Type-A (Fine, larger, expensive) • Type-B (Coarse, small, cheaper) • Reaction Wheel • …. Example of “semi-order-made” approach Um… IRU should be Type-B for my mission… • To achieve this concept, one important thing is to distinguish and separate the AOCS into two categories: • one which is independent of the AOCS components selection, and • ones dependent on components.

13. New Architecture for AOCS Data Interface Conventional ISAS approach For slightly difference component configuration… • Monolithic AOCS computer unit. • Effective to minimize weight, size, number of parts, … • Component change affect wide area in the unit.

14. New Architecture for AOCS Data Interface [Cont’d] New approach • Distributed AOCS computer unit. • Minimize the are influenced by the component change. • Weight etc. are not minimum.

15. New Architecture for AOCS Data Interface [Cont’d] In this novel approach, function to interface each component is distributed in the individual ACIM (attitude control interface module), not implemented in the computer unit. Thus the connection between ACIMs and the computer unit should behaves just like local bus. Then, what is the appropriate data interface for this purpose? “SpaceWire” is assumed to be a solution, because of Speed, Simplicity, Open standard, …

16. Architecture for ground test equipments For software development & test, flight operation simulation, etc. For static closed loop test, etc.

17. Conclusion • A standard bus system for the ISAS/JAXA’s new series of small scientific satellites has been presented. Some ideas and concepts to enhance flexibility of the bus system have been shown. • The TOPS project which is the first satellite of the series will proceed to the Phase B soon. • On and after the second flight, the development time will be no more than two years from determination of mission interfaces to the launch including integration tests. • As the estimated cost is worldwide competitive, the standard bus has potential to be applied to earth observation or disaster monitoring missions in the future. Thank you very much for your attention.