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Two groups of heat transfer problems related tofor the thermal protection of space vehicles can be distinguished. The first group includes complex problems associated with the entry of a space vehicle into the atmosphere of a planet. In this case, the heat flux to the protected surface is the result of the interaction between the radiative and convective heat transfer from the ambient medium, and in which the relative role of these heat transfer modes changes significantly as the space vehicle descends from the rarefied upper atmosphere into the dense layers of the atmosphere. This section is focused on other problems in which external thermal radiation is dominant.

When the a space vehicle moves outside the planetary atmosphere, most of the external heat flux to its surface is associated with thermal radiation. When flying far from the Sun, in addition to direct solar radiation, the visible solar radiation reflected by the planets, as well as the intrinsic infrared radiation of the closely spaced planets, should be taken into account. In this case, the total heat flux to the space vehicle surface is, as a rule, small and for thermal protection it is sufficient to use a relatively light, non-destructive thermal insulation. The features of thermal protection of spacecraft based on multilayer vacuum insulation are considered in the first article of in this section.

Special thermal conditions take place for solar probes designed to study the Sun from a close distance, when the flux of thermal radiation to the surface of the probe is very large and even the use of modern ablative materials may turn out to be insufficient for reliable thermal protection. Therefore, in the second article of in this section, a recently proposed method of partial shielding of solar thermal radiation using a cloud of micron-sized particles of zirconium carbide, formed during the thermal destruction of a special composite material used for thermal protection, is considered.

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