Spacecraft thermal control

The thermal control subsystem can be composed of both passive and active items and works in two ways:

• Protects the equipment from overheating, either by thermal insulation from external heat fluxes (such as the Sun or the planetary infrared and albedo flux), or by proper heat removal from internal sources (such as the heat emitted by the internal electronic equipment).
• Protects the equipment from temperatures that are too low, by thermal insulation from external sinks, by enhanced heat absorption from external sources, or by heat release from internal sources.

Passive thermal control system (PTCS) components include:

• Multi-layer insulation (MLI), which protects the spacecraft from excessive solar or planetary heating, as well as from excessive cooling when exposed to deep space.
• Coatings that change the thermo-optical properties of external surfaces.
• Thermal fillers to improve the thermal coupling at selected interfaces (for instance, on the thermal path between an electronic unit and its radiator).
• Thermal washers to reduce the thermal coupling at selected interfaces.
• Thermal doublers to spread on the radiator surface the heat dissipated by equipment.
• Mirrors (secondary surface mirrors, SSM, or optical solar reflectors, OSR) to improve the heat rejection capability of the external radiators and at the same time to reduce the absorption of external solar fluxes.
• Radioisotope heater units (RHU), used by some planetary and exploratory missions to produce heat for TCS purposes.

Active thermal control system (ATCS) components include:

• Thermostatically controlled resistive electric heaters to keep the equipment temperature above its lower limit during the mission's cold phases.
• Fluid loops to transfer the heat emitted by equipment to the radiators. They can be:
  • single-phase loops, controlled by a pump;
  • two-phase loops, composed of heat pipes (HP), loop heat pipes (LHP) or capillary pumped loops (CPL).
• Louvers (which change the heat rejection capability to space as a function of temperature).
• Thermoelectric coolers.

Parker Solar Probe in thermal testing

• Environment interaction
  • Includes the interaction of the external surfaces of the spacecraft with the environment. Either the surfaces need to be protected from the environment, or there has to be improved interaction. Two main goals of environment interaction are the reduction or increase of absorbed environmental fluxes and reduction or increase of heat losses to the environment.
• Heat collection
  • Includes the removal of dissipated heat from the equipment in which it is created to avoid unwanted increases in the spacecraft's temperature.
• Heat transport
  • Is taking the heat from where it is created to a radiating device.
• Heat rejection
  • The heat collected and transported has to be rejected at an appropriate temperature to a heat sink, which is usually the surrounding space environment. The rejection temperature depends on the amount of heat involved, the temperature to be controlled and the temperature of the environment into which the device radiates the heat.
• Heat provision and storage.
  • Is to maintain a desired temperature level where heat has to be provided and suitable heat storage capability has to be foreseen.

For a spacecraft the main environmental interactions are the energy coming from the Sun and the heat radiated to deep space. Other parameters also influence the thermal control system design such as the spacecraft's altitude, orbit, attitude stabilization, and spacecraft shape. Different types of orbit, such as low Earth orbit and geostationary orbit, also affect the design of the thermal control system.

• Low Earth orbit (LEO)
  • This orbit is frequently used by spacecraft that monitor or measure the characteristics of the Earth and its surrounding environment and by uncrewed and crewed space laboratories, such as EURECA and the International Space Station. The orbit's proximity to the Earth has a great influence on the thermal control system needs, with the Earth's infrared emission and albedo playing a very important role, as well as the relatively short orbital period, less than 2 hours, and long eclipse duration. Small instruments or spacecraft appendages such as solar panels that have low thermal inertia can be seriously affected by this continuously changing environment and may require very specific thermal design solutions.
• Geostationary orbit (GEO)
  • In this 24-hour orbit, the Earth's influence is almost negligible, except for the shadowing during eclipses, which can vary in duration from zero at solstice to a maximum of 1.2 hours at equinox. Long eclipses influence the design of both the spacecraft's insulation and heating systems. The seasonal variations in the direction and intensity of the solar input have a great impact on the design, complicating the heat transport by the need to convey most of the dissipated heat to the radiator in shadow, and the heat-rejection systems via the increased radiator area needed. Almost all telecommunications and many meteorological satellites are in this type of orbit.
• Highly eccentric orbits (HEO)
  • These orbits can have a wide range of apogee and perigee altitudes, depending on the particular mission. Generally, they are used for astronomy observatories, and the TCS design requirements depend on the spacecraft's orbital period, the number and duration of the eclipses, the relative attitude of Earth, Sun and spacecraft, the type of instruments onboard and their individual temperature requirements.
• Deep space and planetary exploration
  • An interplanetary trajectory exposes spacecraft to a wide range of thermal environments more severe than those encountered around Earth's orbits. Interplanetary mission includes many different sub-scenarios depending on the particular celestial body. In general, the common features are a long mission duration and the need to cope with extreme thermal conditions, such as cruises either close to or far away from the Sun (from 1 to 4–5 AU), low orbiting of very cold or very hot celestial bodies, descents through hostile atmospheres, and survival in the extreme (dusty, icy) environments on the surfaces of the bodies visited. The challenge for the TCS is to provide enough heat-rejection capability during the hot operating phases and yet still survive the cold inactive ones. The major problem is often the provision of the power required for that survival phase.

The temperature requirements of the instruments and equipment on board are the main factors in the design of the thermal control system. The goal of the TCS is to keep all the instruments working within their allowable temperature range. All of the electronic instruments on board the spacecraft, such as cameras, data-collection devices, batteries, etc., have a fixed operating temperature range. Keeping these instruments in their optimal operational temperature range is crucial for every mission. Some examples of temperature ranges include

• Batteries, which have a very narrow operating range, typically between −5 and 20 °C.
• Propulsion components, which have a typical range of 5 to 40 °C for safety reasons, however, a wider range is acceptable.
• Cameras, which have a range of −30 to 40 °C.
• Solar arrays, which have a wide operating range of −150 to 100 °C.
• Infrared spectrometers, which have a range of −40 to 60 °C.

• Various composite materials
• Heat rejection through advanced passive radiators
• Spray cooling devices (e.g. liquid droplet radiator)
• Lightweight thermal insulation
• Variable-emittance technologies
• Diamond films
• Advanced thermal control coatings
  • Microsheets
  • Advanced spray on thin films
  • Silvered quartz mirrors
  • Advanced metallized polymer-based films
• 3D printed evaporators for Loop Heat Pipes [3]
• Space Copper-water Heat Pipes for chip-level cooling [4]

A major event in the field of space thermal control is the International Conference on Environmental Systems, organized every year by AIAA. Another is the European Space Thermal Analysis Workshop

Sunshield full-size test for the James Webb Space Telescope

In spacecraft design, a Sun shield restricts or reduces heat caused by sunlight hitting a spacecraft.[5] An example of use of a thermal shield is on the Infrared Space Observatory.[5] The ISO sunshield helped protect the cryostat from sunlight, and it was also covered with solar panels.[6]

Not to be confused with the concept of a global-scale Sun shield in geoengineering, often called a Space sunshade or "Sun shield", in which the spacecraft itself is used to block sunlight to a planet.[7]

An example of a sunshield in spacecraft design is the sunshield on the James Webb Space Telescope.[8]

• Environmental control and life-support system
• Space sunshade
• Temperature control

• Gilmore, D. G., “Satellite Thermal Control Handbook”, The Aerospace Corporation Press, 1994.
• Karam, R. D., Satellite Thermal Control for Systems Engineers, Progress in Astronautics and Aeronautics, AIAA, 1998.
• Gilmore, D. G., “Spacecraft Thermal Control Handbook 2nd ed.”, The Aerospace Corporation Press, 2002.
• De Parolis, M. N., and W. Pinter-Krainer. Current and Future Techniques for Spacecraft Thermal Control 1. Design Drivers and Current Technologies. 1 Aug 1996. Web: 5 Sep 2014.