Temperature control in space

A temperature control system or thermal control system of a satellite or spacecraft is the term used to describe all technical systems and measures for monitoring, controlling and regulating the temperature on board in all phases of the flight.

Since a spacecraft the vacuum of outer space is exposed to a receptacle and especially supply energy in the form of heat to the environment is by heat conduction not possible. This is one of the essential factors that have to be taken into account when designing a satellite, since it means that the temperature (or rather the thermal equilibrium that is established) of a satellite can only be regulated by radiation and radiation (see also gray or black body ) . Careful planning and control are necessary to prevent individual systems and parts of the satellite or the entire satellite from overheating or freezing. Depending on the type of spacecraft (extremely between the sun and deep space probes), extensive measures are sometimes necessary to prevent the absorption or generation of heat or its loss.

Influencing the temperature

In order to influence the temperature of a spacecraft or its systems, it is only possible to change one of the following parameters:

• Heat input:
  • From outside: absorption of radiation from the sun or other nearby celestial bodies,
  • From inside: waste heat from devices on board ( dissipation )
• Heat emission: emission of heat radiation into space
• Heat transfer: heat radiation , convection and heat conduction within the satellite

Control is possible with the help of the following measures:

• passive: about the surface properties ( coating , polishing , roughening ), thermal insulation , heat shield (e.g. gold foils), blinds or panels, thermally conductive materials, surface geometry
• semi-passive: via phase transformation (e.g. heat pipe ), bimetal flakes
• active: via heating elements , cooling circuits , air conditioning systems (especially with manned satellites), alignment of the satellite

The aim of temperature control is to keep the components within the intended temperature range for storage and operation and, if necessary, a suitable temperature for a human crew.

The typical permissible operating temperatures of satellite components differ and are used in chemical processes in engines at 10 to 120 ° C in the tank at 10 to 40 ° C (Einstofftanks) in batteries at -10 to 25 ° C, for electrical components with transponders at 10 to 45 ° C, for earth sensors at −10 to 55 ° and for mechanical components with twist wheels at 0 to 45 ° C and for antennas at −170 to 90 ° C.

Theoretical model of temperature change

In order to avoid that individual components are suddenly no longer in their intended temperature range (overheating, freezing), simulations are carried out. The satellite is divided into so-called nodes (areas assumed to be isothermal) that exchange heat with each other and with the environment. The heat exchange variables are added for each of these nodes.

If the satellite is in thermal equilibrium (i.e. it no longer heats up or cools down and the change in temperature approaches zero), the following applies:

${\displaystyle {\frac {dT}{dt}}=0={\frac {1}{mC_{W}}}\cdot \sum P}$

Here is the power , m is the mass and the heat capacity . ${\displaystyle P}$ ${\displaystyle C_{W}}$

In many cases, these numerical simulations are supplemented or completed by tests. Thermal vacuum chambers, if necessary with solar simulation, are used for this.

The temperature development of the individual components of a satellite are influenced by the following factors near the earth:

Heat absorption

Heat can be absorbed by the following sources:

• Solar radiation :

${\displaystyle P_{\text{Sonne}}=\alpha _{\text{Sat}}\cdot A_{\text{eff}}S\cdot \cos \varphi }$

with degree of absorption (absorbed radiation / total radiation), effective area with regard to the sun , solar constant , relative angle ${\displaystyle \alpha _{\text{Sat}}}$ ${\displaystyle A_{\text{eff}}}$ ${\displaystyle S}$ ${\displaystyle \left(1367\,\mathrm {\tfrac {W}{m^{2}}} \right)}$ ${\displaystyle \varphi }$

• Albedo (sunlight reflected from the surface of a celestial body e.g. the earth):

${\displaystyle P_{\text{Albedo}}=\alpha _{\text{Albedo}}A_{\text{eff}}\cdot S\cdot 0{,}35\cdot \cos \varphi \left({\frac {R_{\text{Erde}}}{R_{\text{Sat}}}}\right)^{2}}$

with degree of absorption , effective area with regard to albedo , solar constant , degree of reflection of the albedo, relative angle , earth radius , orbit radius ${\displaystyle alpha_{\text{Albedo}}}$ ${\displaystyle A_{\text{eff}}}$ ${\displaystyle S}$ ${\displaystyle \varphi }$ ${\displaystyle R_{\text{Erde}}}$ ${\displaystyle R_{\text{Sat}}}$

• Geothermal : ${\displaystyle P_{\text{Erde}}=\varepsilon _{\text{Erde}}\cdot A_{\text{Erde}}\cdot 250{\frac {W}{m^{2}}}\left({\frac {R_{\text{Erde}}}{R_{\text{Sat}}}}\right)^{2}\cos \varphi _{\text{Erde}}}$

with the emissivity of the earth , the area of ​​the earth visible from the satellite , emitted power, angle with respect to the earth , or correspondingly for other celestial bodies. ${\displaystyle \varepsilon _{\text{Erde}}}$ ${\displaystyle A_{\text{Erde}}}$ ${\displaystyle \varphi _{\text{Erde}}}$

• Space radiation:

${\displaystyle P_{\text{Weltraum}}=A\cdot \varepsilon \cdot \sigma \cdot T_{\text{Weltraum}}^{4}\cdot e_{\text{Weltraum}}}$

with emissivity of space , Stefan-Boltzmann constant , temperature of space (mostly negligible) ${\displaystyle \varepsilon }$ ${\displaystyle \sigma }$ ${\displaystyle T_{\text{Weltraum}}=4\,\mathrm {K} }$

• Dissipation : ${\displaystyle P_{\text{Disp}}}$
• Aerodynamic heating: In the case of orbits very close to the earth or planets, energy absorption through friction with the earth's atmosphere (aerodynamic heating) is also possible and must be taken into account.

Heat transfer to other components

Heat transfer to other parts of a spacecraft can occur through:

• convection
• Conduction
• Thermal radiation

Heat release into space

The only way to give off heat in space is to radiate it into space:

${\displaystyle P_{\text{Abstrahlung}}=-A\cdot \varepsilon \cdot \sigma \cdot T^{4}}$

Usually special radiators are used for this , which must not be exposed to sunlight for correct functioning. For missions close to the sun it may be necessary to coordinate the entire satellite design with the design of the radiating surfaces.

literature

• Ernst Messerschmid and Stefanos Fasoulas; Space systems: an introduction with exercises and solutions ; ISBN 3540776990
• Ley, Wittmann, Hallmann; Space Technology Handbook; ISBN 9783446411852

Web links

• ISS Active Thermal Control System (Boeing) (PDF; 1.3 MB)
• Thermal Control System of the French Spot Satellite