The International Space Station 160 to 120V

The International Space Station is truly the largest and most complex space structure ever built, with 16 international partners. The complete assembly shown in Figure 4.15 weighs one million pounds and the total interior space of its six laboratories equals two Boeing 747 aircraft. It is taller than a 30-story building and wider than the length of a football field. The attitude is controlled within 1° stabilized at a rate less than 0.1 °/s. It is in 335 to 500-km low Earth, 51.6° inclined orbit with a 90-min orbit period and 35-min eclipse. The experiments planned for its on-board laboratories are targeted to enhance our understanding in:

• The Earth, by viewing 75% of its surface

• Long term exposure of micro-gravity on humans

• Producing new medicines and materials

Like any other large space programs, such as the manned landing on Moon, the engineering and technology spin-offs on our everyday lives on Earth are the side benefits.

The ISS power system generates 105 kW using a solar array with an area of nearly 1 acre. The loads are powered during an eclipse by thirty-eight

NASA artist's conception

FIGURE 4.15 International Space Station with solar array and other modules in view. (Source: NASA.)

NASA artist's conception

FIGURE 4.15 International Space Station with solar array and other modules in view. (Source: NASA.)

81-Ah NiH2 batteries3 via bi-directional battery charge-discharge units (BDCUs), autonomously controlled by the bus voltage set point. It has two interconnected power systems; the 160/120-V US built system, and 120/ 28-V dual voltage Russian built system to power the US, European, and Japanese modules. The two systems are generally independent, but are interconnected via d.c. converters to allow bi-directional transfer of power. Figure 4.16 is the functional block diagram of the US system, including the converter at the interface with the Russian power source. The US system is described below.3

The solar array is made of four modules (wings) for a total of 76 kW power under nominal conditions, and more during favorable conditions. Each wing consists of two thin blankets held under tension on each side of a central collapsible mast. The entire assembly turns on a ft gimbal, which provides one axis of rotation for sun pointing. A second orthogonal axis rotation is provided at the a gimbal, where the entire solar array connects to the rest of the truss structure of the station. The 76 kW from the US system added to the 29 kW from the Russian system makes the 105 kW capability of the complete station assembly. Each of the four solar modules is made in 82 strings of eighty 8 x 8 cm PV cells and coverglass for protection against space charged particles. The PV cells are crystalline silicon with 14.5% EOL efficiency over a 15-year life. The module printed circuit is Kapton/copper/ Kapton laminate welded to each cell to provide a series interconnection for

Solar array

Primary distribution

138-173 Vdc 48^Wg?mbal ssu -QD-

126-173 Vdc

21=9kW Alpha ¿1-aKW gimbal

DCSU

Solar array

DDCU

BCDU

Secondary distribution 120-126 Vdc

19.1 kW

DDCU

ARCU

BATS

ARCU

PFCS

ES EC

Russian sources (1)

loads

(1) See fig. 5 for corresponding connection points and FGB architecture

Sunlit period (charging) 54.9 (minimum) ^^ Eclipse period (discharging) 36.5 (maximum)

ARCU - American-to Russian Converter Unit BATS - Batteries

BCDU - Battery Charge/Discharge Unit DCSU - DC Switching Unit MBSU - Main Bus Switching Unit RACU - Russian-to-American Converter Unit RPCM- Remote Power Controller Module SSU - Sequential Shunt Unit

FIGURE 4.16 International Space Station single channel power flow diagram. (From E.B. Gietl et al. NASA Report No. 210209, 2000.)

the assembly. Parallel bypass diodes are used every eight cells for reliability in case of cell damage and to avoid reverse current heating during prolonged shadows. Sequential shunt units (SSUs) on the solar arrays operate at 20-kHz switching frequency.

The seasonal sun-pointing is done by ft gimbals and the orbit sun-following by a drive and roll rings. For currents of the space station magnitudes, the roll rings provide superior power transfer performance over the slip rings with rubbing contacts, as described further in Chapter 22.

The solar array output voltage is 160 V, which is the highest voltage that can be practically used in a low Earth (high plasma) orbit in view of potential plasma arcing and/or leakage current concerns. The 160 V is stepped down to 120 V using d.c.-d.c. converter units (DDCUs), each rated at 4.25 kW, for utilization inside the user modules. The DDCUs provide 150% current limiting capability and 20 dB isolation between the generation point and the distribution points for personnel safety.

The solar array area and the operating voltage are greater than in any other spacecraft flown before. Therefore, the nature of the single point ground on the ISS in high plasma in LEO poses an arcing problem. To preclude such arcing, a device called the plasma contactor located on the truss creates a plume of ionized xenon gas, which acts as a conductive bridge between the station and the plasma. It protects the array and other conductive surfaces of the station from arcing, pitting, and erosion.

The battery is made of 48 battery packs, each with thirty-eight 81-Ah IPV NiH2 cells, designed for 5 years (40,000 charge/discharge cycles) at 35% depth of discharge (DOD), although the actual operational DOD is around 15%. With two packs connected in series, the battery voltage varies between 95 and 115 V. Battery replacements are planned every 5 years to assure the power margin. With 15 years life of the station, the battery replacements every 5 years constitute significant recurring cost. The maximum charge rate is 50 A, which tapers down to 40, 27, 10 and 5 A, and finally to 1 A trickle charge rate. No bypass diodes are used in either direction, but the cells are closely matched. An open cell will lose the entire battery string. The operating temperature is in the range of 0 to 10 °C.

The switching and fault protection is achieved by solid-state remote power controllers (RPCs) in six ratings from 3.5 to 65 A in both current limiting and non-current limiting designs. The RPCs trip at different set points of over-current, over-voltage, and under-voltage to isolate faults as close as possible to the faulted equipment. The solid state power controllers (SSPCs) provide switching and protection. They reset automatically for critical loads and remain tripped for non-critical loads. The upstream coordinated fault protection in the 160-V segment is achieved by larger main bus switching units (MBSUs) containing a remote bus isolator (RBI). The RBIs are essentially large relays capable of interrupting up to 350 A of d.c. fault current.

The power system stability is also a serious concern because the loads on the station are constantly changing as new scientific experiments are brought on board. As a result, the output impedance of the DDCUs and the input characteristic of the loads have been specified to ensure stability with any combination of the expected loads. The stability criteria under such diverse load combinations are discussed in Chapter 14.

A hierarchy of redundant computers linked via the MIL-STD-1553 bus controls the ISS power system. The computers autonomously control many functions such as sun tracking, battery energy storage, and thermal control. The power flow balance between major segments of the station is coordinated by the on-board command and control systems, which also provide interface control between the segments.

Future plans for new technology developments on the ISS that are related to the electrical power systems include: (1) replacing the silicon solar cells with GaAs panels to increase conversion efficiency, (2) using high discharge cycle NiH2 batteries, and (3) using flywheel energy storage to significantly increase the specific energy.

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