Distributed Power System For Large Orbital Infrastructures: Architecture, Requirements Analyses and Lesson Learned Antonio Ciccolella ESA-ESTEC, D-TEC/EEE, Keplerlaan 1 - 2201 AZ Noordwijk, The Netherlands Abstract learned during the development and operational phase of Modules and payloads. Distributed power systems offer many benefits to system designers over central power systems such as I. INTRODUCTION: ISS POWER ARCHITECTURE reduced weight and size. Distributed systems also allow the designers to control the quality of power at The ISS is a joint project of six space agencies: the different loads and subsystems, since DC-DC U.S. National Aeronautics and Space Administration converters allow close regulation of output voltage (NASA), Russian Federal Space Agency, Japan under wide variations of input voltages and loads. Aerospace Exploration Agency (JAXA), Canadian Distributed power systems also provide a high degree Space Agency (CSA/ASC), Brazilian Space Agency of reliability because of the isolation provided by (AEB) and the European Space Agency (ESA). DC/DC converters; it is very easy to isolate system The space station is located in orbit around the Earth failures and provide redundancy. These systems are at a nominal altitude of approximately 360 km (220 also very flexible and easily expanded. miles), a type of orbit usually termed low Earth orbit. The actual height varies over time by several This paper addresses the DC distributed power kilometers due to atmospheric drag and re-boosts. It system of the International Space Station, which is a orbits Earth at a period of about 92 minutes. specific case of this kind of distributed system. In many ways the ISS represents a merger of previously planned independent space stations: It is a channelized, load following, DC network of Russia's Mir 2, United States' Space Station Freedom solar arrays, batteries, power converters, switches and and the planned European Columbus Attached cables which route current to all user loads on the Pressurized Module (APM). Today it represents a station. The completed architecture consists of both permanent human presence in space, as it has been the 120-V American and 28-V Russian electrical manned with a crew of at least two since November networks, which are capable of exchanging power 2, 2000. through dedicated isolating converters. It is serviced primarily by the Space Shuttle, Soyuz and Progress spacecraft units. In 2006 a new The presence of DC/DC converters required special European service vehicle, the Automated Transfer attention on the electrical stability of the system and Vehicle (ATV), will be operational. ISS is still being in particular, the individual loads in the system. This built, but it is home to some experimentation already. was complicated by complex sources and undefined At present, the station has a capacity for a crew of loads with interfaces to both sources and loads being three. designed in different countries (US, Russia, Japan, At its final integration, the ISS characteristics will be: Canada, Europe, etc.). These issues, coupled with the program goal of limiting costs, have proven to be a • Dimensions: 87 m (L) x 106 m (W) significant challenge to the program. • Weight: 425 000 Kg • Pressurised Volume: 1160 m3 As a result, the program used an impedance • Power: 110 kW (total), 45 kW for Users. specification approach for system stability. This approach is based on the significant relationship NASA is responsible for the overall system, between source and load impedances and the effect including Space Station integration and verification, of this relationship on system stability. It is limited in supported by proactive and intense interactions with its applicability by the theoretical and practical limits all the participating parties. on component designs as presented by each system The ultimate power source for the station is the segment. Consequently, the overall approach to sunlight, which is incident on solar arrays deployed system stability implemented by the ISS program as paired sets. consists of specific hardware requirements coupled Each array wing consists of two thin blankets held with extensive system analysis and hardware testing. under tension on each side of a central collapsible Highlights of both experimental and analytical mast. The entire assembly turns on a “beta gimbal”, activities will be shown, as well as some lesson which provides one axis of rotation for solar pointing. A second orthogonal axis of rotation is provided at power directly to core system loads and to all the “alpha gimbal”, where the entire solar power payloads. module connects to the rest of the truss structure of the station (see Fig. 1). These arrays each provide a The DDCUs provide excellent isolation between total of 25 kW of power during the sunlit portion of input and output ports, ensuring that perturbations on the 90-minute, low-earth orbit of the station. Due to one branch of the primary system will have minimal solar cell degradation over time, and different impact on other branches. Downstream from each pointing angles, the optimum output voltage will vary DDCU are Remote Power Controller Modules between 160 and 140 Volts over the life of the (RPCMs), which are computer-controlled, solid-state station. circuit breakers that branch out the power to the loads. II. DESIGN ISSUE: IMPEDANCE MATCHING The Power System of the ISS is distributed rather than centralised, i.e. the power is processed by multiple cascaded stages of switching regulators. The canonical arrangement of distributed power systems consists mainly of: • A stage of regulators, acting as line conditioners, which take the unregulated input voltage and convert it to a regulated bus voltage. This bus distributes power to the system. • Cascaded stages of parallel regulators, Fig.1 – Power Architecture of ISS (one channel) acting as load converters, which take the power from the above bus and generate the These are remotely controlled switching boxes, used appropriate output voltage required by the to route power between redundant channels, and to loads. direct power to Dc-to-Dc Converter Units (DDCUs) and local loads. The adoption of a Distributed Power System for the ISS is dictated, inter alia, by the difficulty in The first switching converters in the power string are maintaining regulation at the load location with a the BCDUs, which regulate charging of the Nickel- centralised power distribution, due to the possible Hydrogen batteries during sun exposure, and serve long distance between load and source. By placing the dual role of discharge control and bus regulation intermediate converters closer to the loads, the during eclipse. The units are bidirectional, stepping requested voltage is regulated with higher accuracy. the voltage down to the battery level of Another advantage of the Distributed Power System approximately 72 Volts when charging, and stepping for large spacecraft is that cascading regulators the battery voltage up to the main bus voltage when inherently enhance isolation between source and discharging. loads. Furthermore, the use of parallel load regulators implies modularity, which allows the spacecraft Switching between charge and discharge is design to flexibly manage power reconfiguration and automatically triggered by the bus voltage set-point, reliability by adding redundant modules. even during sunlight, if load demand exceeds the ability of the solar arrays to provide full power, the These features of the distributed power system are batteries will switch in to pick up the slack. achieved at the expense of higher system complexity, which may render the analysis and the design of the ARCUs allow interconnection between the US-built system as impracticable if it is considered as a whole. 160/120-Volt system and the Russian-built 28-Volt However, the intrinsic modularity suggests the use of system. Several of the Russian modules have their an approach for both the design and the analysis, own independent solar arrays, providing important which foresees the complete system as partitioned in levels of redundancy for power availability. When smaller power subsystems that can be independently power is being transferred between systems these and individually considered. converters play a role in the overall bus impedance of both the systems. The system is then built through the appropriate The DDCUs provide the interface between the interconnection of the previously mentioned primary 160-Volt system and the more tightly subsystems, whose interaction needs to be carefully regulated 120-Volt secondary system, which provides analysed eventually to generate compatible their input filter loaded by a negative resistance, specifications. whose magnitude is |R|=V2/P. This is not valid when the input filter is integral part This approach, widely used in the frame of the ISS of the power cell, which can occur for some [1], is based on the underlying assumption that the topologies. The above impedance matching concept interacting subsystems are linear, thus small signal was applied to determine realistic boundaries to levy stability is the objective to achieve. requirements on loads and subsystems in flight elements (i.e. APM, JEM). Namely, we consider (Fig. 2) a source block with It implied analysis of multiple loads with a known forward gains FS and a load block with forward gain source and the requirement were responsibility of FL. We designate ZS as the source impedance and ZL NASA, with the necessary attribute of being not as the load impedance. overly conservative or impractical to implement NASA initially identify two load’s category: • Complex (load assemblies or equipment racks): interface B • Simple (individual loads): interface C Interface B category (required to have their own power distribution and protection) was subdivided in function of their feeder rating (10-12 A, 25-30 A, 50- 60 A). Fig.2 - Source Block cascaded with a Load Block Interface C category was subdivided in function of their line branch rating (1.5-3.5 A, 10-12 A, 25-30 A) Working with the hybrid g-parameters one can verify Four ranges of harness length were also considered.