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Redundant unit, 250 or 1,000 bits/s 64 word, 8 bit frame 5 subcommutated channels

required temperatures. The sources of heat include solar radiation, Earth-reflection and infrared radiation, and electrical energy dissipated in the electrical components. Conventional electronics operate at temperatures close to room temperature (25 cC) and will tolerate temperature variations of about ±20 cC. Battery cells, particularly nickel-cadmium cells, are more sensitive to temperature than most electronics. But they can still stand temperature ranges of 5 °C to 20 "C. We can control the temperatures of compartments for conventional electronics by coating or insulating then- outer surfaces. We select these coatings to strike a balance between the heat absorbed and the heat radiated to space. The coatings include various paints and tapes, and second surface glass mirrors. The weight of such coatings is almost independent of the quantity of heat dissipated and seldom exceeds 4% of the spacecraft dry weight The thermal coatings, particularly insulation, can close out compartment openings and may also shield components from electromagnetic radiation.

Components which have stringent temperature requirements or which dissipate large amounts of electrical power require more extensive thermal control. For example, we usually place gyros and precision oscillators in insulated compartments, or ovens, with active electrical heaters to control temperatures carefully. Traveling wave tubes and other elements which dissipate a lot of power concentrate their dissipation locally and produce hot spots. Normally, we conduct heat away from such hot spots and spread it over a thermal panel where it radiates to space. If the hot spots are less than 50 W, we can simply increase the thickness of the mounting panel so it will cany the heat away. But more intense hot spots and equipment that must meet tight thermal limits usually require heat pipes to move the heat and thus equalize temperatures.

The process of thermal design for spacecraft proceeds as follows. The first step is to identify heat sources and the location of radiating panels to dispose of excess heat The heat sources include internal dissipation and external radiation from the Sun and the Earth. Often, particular surfaces of the spacecraft are not exposed or only partially exposed to the Sun or Earth. These faces are preferred locations for radiating panels and for components which dissipate large amounts of heat The latter should be next to the radiators. We can compute the total amount of radiating area from the static-heat-balance equation for the entire spacecraft But a component's thermal performance may differ markedly from the average.

Most power for thermal control goes to heaters that keep components from getting too cold. Heaters also compensate for imperfections in insulation or heat leaks, and are used to heat areas such as articulation joints that are difficult to insulate and dissipate little heat internally. Heaters or heater-controlled heat pipes (see Sec. 11.5) offer very tight control or control at a particular temperature value for certain components. A typical medium-sized spacecraft (1,000 W) consumes 20 W in the thermal subsystem plus any power required for special thermal control. In most cases, heaters can operate from primary power.

10.4.6 Power Subsystem

The power subsystem generates power, conditions and regulates it stores it for periods of peak demand or eclipse operation, and distributes it throughout the spacecraft The power subsystem may also need to convert and regulate voltage levels or supply multiple voltage levels. It frequently switches equipment on or off and, for increased reliability, protects against short circuits and isolates faults. Subsystem design is also influenced by space radiation, which degrades the performance of solar cells. Finally, battery life often limits the spacecraft's lifetime.

Earlier in this chapter, I described how to prepare a power budget for the spacecraft. This budget includes most of the information we need to size the power subsystem: the spacecraft's needs for operating power, storage requirements, and how the power subsystem degrades over the spacecraft's lifetime. The remaining steps to size the power subsystem are selecting a solar-array approach, sizing the array, sizing the batteries and the components that control charging, and sizing the equipment for distributing and converting power.

Solar arrays are generally planar, cylindrical, or omnidirectional. Planar arrays are flat panels pointed toward the Sun. Their power output is proportional to the projection of their area toward the incident sunlight Three-axis-stabilized spacecraft normally use planar arrays. Cylindrical arrays appear on spin-stabilized systems in which the spin axis is perpendicular or nearly perpendicular to the Sun line. The output of a solar-cell array is nearly proportional to the amount of solar energy intercepted, and the projected area of a cylinder is 1/« times the total area. Thus, the cylindrical array should have approximately ft times as many cells as a planar array with the same power rating. But temperature effects slightly favor the cylindrical array, so the actual ratio is closer to 1/2.5. If the spacecraft can receive sunlight from any aspect then its array must have equal projected area in all directions. In other words, it must have an omnidirectional array. A sphere has this property, but paddles or cylinders combined with planar panels are also possible. The total area of an omnidirectional array must be approximately 4 times the projected area, so an omnidirectional array has about 4 times the area of a planar array with the same power rating.

The required area of a planar solar array is related to the required power, P, the solar constant (1,367 W/m2), and the conversion efficiency of the solar-cell system. Although cells have had efficiencies as high as 30%, practical array designs range from 5% to 15% when taking into account the operating conditions and degradation at end-of-life. An array with an efficiency of 7% would have a required area of p

where Aa is in m2 and P is in watts. This area is characteristic of current arrays. The mass of a planar array with specific performance of 25 W/kg is

where Ma is in kg and P in watts. Current designs range from 14 to 47 W/kg at end-of- life. The high end would provide 66 W/kg at beginning-of-life. Solar arrays mounted on the spacecraft's body usually weigh less than planar arrays.

Rechargeable nickel-cadmium or nickel-hydrogen batteries are the usual devices for energy storage for unmanned spacecraft They are available in various sizes and are highly reliable even though their performance characteristics are quite complex. The battery often represents one of the most massive components in the spacecraft It also is very sensitive to temperature and to the use profile. Nickel-cadmium, and to a lesser extent nickel-hydrogen batteries perform best when operated between 5 °C and 20 °C. This range is both lower and more restricted than the temperature requirements for most electronic components. The battery also has complex wear-out mechanisms, thus limiting cycle life as a function of depth-of-discharge. Other variables—temperature, rate of charge, rate of discharge, and degree of overcharge—also affect cycle life but in a less well-defined way. If a battery has shallow discharge cycles, it loses capacity. To counter this tendency, most spacecraft recondition their batteries from time to time by discharging them completely.

We determine a battery's capacity from the energy it must produce (discharge power times discharge duration) and from its depth-of-discharge. We select the battery's depth-of-discharge to meet cycle life requirements. Table 10-26 gives guidelines on depths-of-discharge for nickel-cadmium and nickel-hydrogen batteries. Section 11.4 discusses these concepts in more detail.

TABLE 10-26. Allowed Battery Depth-of-Dlscharge vs. Cycle Life.

Cycle Life

Battery Type

Depth of Discharge

Less than 1,000 cycles

NiCd

80%

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