devices or when required temperature tolerances are very tight Once we thoroughly understand the thermal control challenges, we enter Step 4 of the process to develop thermal designs that can be used to satisfy our requirements. These may range from applying different paints, multi-layer insulation blankets, and component placement, to the use of more sophisticated devices like cryogenic cooling systems, heat pipes, and thermostatically-controlled heaters.
Once we have identified a potential thermal control approach and configuration, we proceed to Step 5 where we determine the radiator and heater requirements for the spacecraft and its components. We consider two worst-case conditions: worst-case hot, where the spacecraft is in the Sun and maximum power is being dissipated, and worst-case cold, where the spacecraft is in eclipse and dissipating minimum power. During this step we also try to understand the performance of the thermal control system over time taking into account degradation of thermal control surfaces and extraordinary thermal events or circumstances.
In Step 6 we use the information generated previously to estimate the mass and power of the thermal control system. As usual we document the results (in Step 7) and repeat the entire process until we create a thermal control system that meets the requirements and constraints at an acceptable mass, cost, and risk.
Each of the steps in this process is discussed in more detail below. For a much more extensive discussion of the thermal control process for space systems see Gilmore  or Karam .
Spacecraft thermal control is a process of energy management in which the thermal environment plays a major role. Over the course of the development and operational life cycle, a spacecraft will be exposed to environments encountered during ground testing, transportation, launch, orbit transfer, and operational orbits with nominal and safehold attitudes. During ground operations, convection with ambient air and radiant heat exchange with surrounding objects are the principal environmental influences. During launch ascent, radiant heating from the inside surfaces of the booster fairing and, after the fairing is jettisoned, free-molecular heating due to friction with the atmosphere are the dominant environmental drivers. Once above the upper reaches of the atmosphere, direct sunlight, sunlight reflected off of Earth or other planets (albedo), and infrared (IR) energy emitted from a planet's atmosphere or surface are the only significant sources of environmental heat. In most cases, the thermal control system is designed to maintain all spacecraft components within allowable temperature limits in the environments encountered on-orbit, while compatibility with ground operations and launch ascent conditions is assured by controlling the environment or limiting the spacecraft's exposure to it
As illustrated in Fig. 11-IS, the overall thermal control of a spacecraft on orbit is usually achieved by balancing the heat emitted by the spacecraft as IR radiation against the heat dissipated by its internal components plus the heat absorbed from the environment; atmospheric convection is absent in space. Because a generic thermal control system capable of maintaining spacecraft temperatures in all environments would be prohibitively heavy and expensive, it is generally more cost effective and practical to custom-tailor a thermal design to each spacecraft and its mission environment This means that the thermal design analysis must consider the worst case hot and cold combinations of waste heat generated by spacecraft components in their various operating modes and the variable environmental heat loads on the spacecraft.
Sunlight is the major source of environmental heating on most spacecraft Fortunately, the Sun is a very stable energy source which is constant to within a fraction of a percent. However, because the Earth's orbit is elliptical, the intensity of sunlight reaching Earth varies approximately ± 3.5%, depending on Earth's distance from the Sun. At summer solstice, Earth is farthest from the Sun, and the intensity is at a minimum value of 1322 W/m2; at winter solstice, the intensity is at its maximum value of 1414 W/m2. The intensity of sunlight at Earth's mean distance from the Sun (1 AU) is known as the solar constant and is equal to 1367 W/m2.
Solar intensity also varies as a function of wavelength, as shown in Fig. 11-15. The energy distribution is approximately 7% ultraviolet, 46% visible, and 47% near (short-wavelength) IR. However, the IR energy emitted by the Sun is of a much shorter wavelength than that emitted by a body near room temperature. This distinction allows for the selection of thermal-control finishes that are very reflective in the solar spectrum but highly emissive to room temperature (long-wavelength) IR, as shown in Fig. 11-15. These finishes, which will be discussed in more detail later, minimize solar heat loads while maximizing a spacecraft's ability to reject waste heat
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