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Spacecraft thermal control subsystems (TCS) essential for both the physical integrity of the satellite and for its efficient operation

Spacecraft are man-made machines that operate in space. An orbiting spacecraft is normally referred to as a satellite, although it is manmade as opposed to a natural satellite like our moon. A spacecraft is typically subdivided into two major parts, the payload and the bus. Where the mission can be defined as the purpose of the spacecraft and is usually identified as the payload part of the spacecraft (e.g. scientific instruments, communications). The thermal control subsystem and other subsystems (e.g. attitude control, electrical power, structures) are part of the bus.


Just as human body has a temperature range that it feels comfortable say 70°F to 72°F, most spacecraft components also have a range
which must be maintained in order to protect the reliability of the components. Thermal control is what allows maintaining the satellite‘s (or another space system) temperatures within set parameters during its lifetime. Thermal control is also what keeps the specified temperature stability for delicate electronics or optical components so as to ensure that they perform as efficiently as possible. Controlling the level of temperature of equipment, payloads, satellites and launchers is essential during all phases of a space mission to protect flight hardware and to guarantee the optimum performance and success of the mission.


The thermal control subsystem is one of the most visually distinctive elements of a space system, composed of distinctive foil-like insulation blankets and white-painted or mirror-like radiators.


Space Thermal Environment

Thermal control issues can have serious impacts to the spacecraft’s mission. Knowledge of both external and internal heat sources must be included in design considerations in order to maintain thermal balance.

The primary sources of externally generated heat for earth-orbiting spacecraft are direct sun, reflected sun from the earth, and infrared (IR) energy emitted from the earth. The spacecraft operating environment largely depends on its orbit type, which is primarily driven by its intended mission. Other parameters also influence the thermal control system design such as the spacecraft’s altitude, orbit, attitude stabilization, and spacecraft shape.


The three common earth-orbiting satellite types are geosynchronous orbit (GEO), highly elliptical orbit (HEO), and low earth orbit (LEO).  of the thermal control system. LEO satellites have an altitude of less than one thousand miles. Satellites in HEO can have a wide range of altitudes for perigee (lowest altitude) to apogee (highest altitude). GEO satellites, orbiting about the equator at some small angle (inclination angle), have an altitude of ≈ 23,000 miles above earth. At this altitude, satellites have the same period of rotation as the earth, appearing fixed relative to earth.


LEO 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.


GEO satellites, orbiting about the equator at some small angle (inclination angle), have an altitude of ≈ 23,000 miles above earth. 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.

Satellites in HEO 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.


For direct sun, the amount of time the sun is “seen” per orbit by the spacecraft also affects its temperature control abilities, especially when extended for long periods of time. Each of the orbital types have a cyclical characteristic where the spacecraft goes from light (i.e.
the sun is present) to dark (i.e. cannot “see” the sun). Sometimes, the spacecraft may be in sunlight or darkness for extended periods of time; worst cases are known as 100% sun or minimum sun respectively. Percent sun refers to the amount of sun present during each orbit.

An interplanetary trajectory exposes spacecraft to a wide range of thermal environments more severe than those encountered around Earth’s orbits. The 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.


For interplanetary space missions, the primary source of external thermal heat is direct sun, secondary is planetary reflected and IR when passing near a planet. This means that most of the time, for interplanetary missions, the heat source intensity will be a function of the spacecraft distance to the sun. Therefore, missions to distant planets (e.g. Jupiter) and further incur very low solar radiances creating an extremely cold spacecraft environment. 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.


Temperature requirements of Subsystems

Thermal control is essential to guarantee the optimal performance and success of the mission because if a component is subjected to temperatures that are too high or too low, it could be damaged or its performance could be severely affected. Thermal control is also necessary to keep specific components (such as optical sensors, atomic clocks, etc.) within a specified temperature stability requirement, to ensure that they perform as efficiently as possible. Large temperature differences within the satellite are also undesirable because they can lead to thermal expansion or contraction, potentially distorting its structure and thereby result in e.g misalignments of optical systems.


It must cope with the external environment, which can vary in a wide range as the spacecraft is exposed to deep space or to solar or planetary flux, and with ejecting to space the internal heat generated by the operation of the spacecraft itself. Thermal control for space applications covers a very wide temperature range, from the cryogenic level (down to -270 deg C) to high-temperature thermal protection systems (more than 2000 deg C).


Thermal control is absolutely essential for both the physical integrity of the satellite and for its efficient operation because electronic equipment have their optimum performance within a certain temperature range. The satellite’s payload will dictate its operating range. Some instruments with infrared detectors for example require extreme low temperatures. Many components have their lifetimes reduced by sustained high temperatures.


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.


Thermal Control System

In spacecraft design, the function of the thermal control system (TCS) is to keep all the spacecraft’s component systems within acceptable temperature ranges during all mission phases. For instance, a piece of equipment could, if encountering a temperature level that is too high, be damaged or its performance could be severely affected. If a component fails sooner than expected, from excessive overheating for example, the spacecraft’s mission could also be shortened.  In space it would hardly be possible to correct such a problem and this is why space thermal control systems – like other space subsystems – need to be properly designed and tested and need to be very efficient and highly reliable.


The thermal control subsystem seeks to maintain the overall temperature to an acceptable level but also to obtain the most adequate temperature distribution within the satellite. It is the task of the thermal engineer to manage the distribution of heat within the satellite so as to ensure that the temperature level is adequate for all phases of a mission (launch, transfer orbit, operation in orbit).

  • 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.


It 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 the desired temperature level where heat has to be provided and suitable heat storage capability has to be foreseen.


Active or passive systems

The level of temperature of a spacecraft is dictated by the balance prevailing between incoming external solar, albedo and planet heat fluxes, heat which is produced internally e.g by electronic equipment and the heat which is rejected to deep space. Just like on earth, heat transfer for a spacecraft is also governed by the following three fundamental ways for the transference of heat energy: convection, conduction, or radiation. Unlike earth however, air cannot be used as a medium to transfer heat in the zero gravity, vacuum of space.


Temperatures are regulated throughout a spacecraft with passive and/or active thermal management techniques. While traditional thermal control techniques have been well demonstrated on large spacecraft, these existing techniques sometimes require additional development for application for small spacecraft applications. Given the increased interest in small spacecraft over the last decade, “miniaturized” applications of these thermal management methods was advanced to ensure adequate thermal control techniques are available for SmallSats.


Passive thermal control system (PTCS) components include:

Passive thermal control requires no input power for thermal regulation of a spacecraft. This can be achieved using several methods and is highly advantageous to spacecraft designers, especially for the CubeSat form factor, as passive thermal control systems are associated with low cost, volume, weight, and risk, and due to their simplicity have been shown to be highly reliable. Ideally, you want all thermal control devices to be passive in order to eliminate the need to utilize part of the electrical subsystem’s power budget. This is because the primary goal of the electrical power subsystem is to maximize the amount of power to the payload.


  • The integration of Multi-Layer Insulation (MLI), thermal coatings/surface finishes, heat pipes, sunshades, thermal straps and louvers are some examples of passive methods to achieve thermal control in a spacecraft. 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. A large proportion of the sun heat flux is blocked with insulation devices called multi-layered insulation blankets (MLIs). The heat is rejected from the satellite to space (which is very cold, at a temperature of about -270 deg C) via radiators.
  • 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 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.


Other “passive” methods include spacecraft design methodologies that help manage thermal loads. These can include structural and electrical design elements that help improve heat transfer via conduction, thereby reducing (or maintaining) component temperatures.

Examples of these include:

thermally isolated structural joints where multiple washers with low thermal conductivity are stacked between fasteners and joined surfaces to limit heat transfer via conduction in specific places and; Thermal washers to reduce the thermal coupling at selected interfaces.

circuit board designs that include copper layers connected by vias that conduct heat away from electrical components through the boards to their connectors/structural connection points (using the thermal mass of the structural bus).

Optical Solar Reflectors

An optical solar reflector (OSR) is used for the thermal control of spacecraft on the sun-facing sides of satellites by reflecting incoming solar radiation while simultaneously radiating internally-generated heat.


In 2018,  team that includes researchers from the University of Southampton (Southampton, England)  developed new technology that could significantly improve spacecraft or satellite exploration. Metamaterial Optical Solar Reflectors (meta-OSRs) are the first-surface coatings on the outside of a spacecraft, designed to effectively radiate infrared heat away from the surface while reflecting most of the optical solar spectrum.


For a satellite or spacecraft, the OSRs play a crucial role in the system’s thermal control. Glued to the external skin of the radiator panels, OSRs are designed to reject solar radiation and dissipate the heat that is generated on board. OSRs are commonly made of quartz tiles that combine thermo-optical properties with an ability to withstand the environment in space. Unfortunately, quartz tiles are heavy and fragile, add significantly to assembly and launch costs, and cannot be applied to curved surfaces. Other commercial solutions based on polymer foils suffer from fast performance degradation and are therefore unfit for missions lasting more than three to five years.


The team demonstrated that a new meta-OSR coating is enabled by the use of metal oxide, a material commonly used for transparent electrical contacts, which, in this instance, is patterned into a metamaterial with very strong infrared emissivity while retaining a low absorption of the solar spectrum. The team also demonstrated a ‘smart’ radiator based on their metamaterial design that allows tuning of the radiative cooling of the spacecraft using another type of metal oxide.


University of Southampton professor Otto Muskens, principal investigator of the study, said, “The meta-OSR technology is entirely based on durable and space-approved inorganic coatings, which can be applied onto flexible thin-film substances with the potential to be developed as a new technology solution. Since the assembly and launch costs of OSRs is several tens of thousands of US dollars per square metre, even small improvements in weight reduction can make a significant change to the space industry.”



In a vacuum, heat is transferred by two means: radiation and conduction. The internal environment of a fully enclosed small satellite is usually dominated by conductive heat transfer, while heat transfer to/from the outside environment is driven via thermal radiation. Thermal radiation heat transfer is controlled by using materials that have certain specific radiative properties, namely: solar absorptivity (implying wavelengths in the range of ~0.3 – 3 µm) and, IR (infrared) emissivity (~3 – 50 µm).


Solar absorptivity governs how much of the incident solar flux a spacecraft absorbs, while IR emissivity determines how well a spacecraft emits its thermal energy to space, relative to a perfect blackbody emitter, and what fraction of thermal radiation from IR sources (e.g., the Earth, Moon) are absorbed by that spacecraft surface. These properties are optical surface properties of a material and can be modified by adding specialized coatings, surface finishes, or adhesive tapes with their own specific coatings.


Thermal control surfaces can be any surface that is used for spacecraft thermal control to include coatings, paints, and finishes. Most internal and external spacecraft components have a thermal control surface to help control its emittance and/or absorptance properties. The performance of these surfaces are characterized by the ratio of absorptivity to emissivity, α / ε, which are characteristics of heat transfer by radiation. For example, white paint has a low ratio and therefore is used as a heat emitter. Ratios greater than 1.0, like blank paint, will get hot when exposed to sunlight.



Coatings are the simplest and least expensive of the TCS techniques. A coating may be paint or a more sophisticated chemical applied to the surfaces of the spacecraft to lower or increase heat transfer. The characteristics of the type of coating depends on their absorptivity, emissivity, transparency, and reflectivity.


Black and white paints are the most common color. Most paints have a high emittance, with varying absorptance and electrical conductivity properties. Black paints have the following characteristic: α ≈ ε ≈ 1.0, which is close to thermal equilibrium. This means that most of the
heat that is absorbed is then emitted. This makes black paint an effectively neutral color, thermally. Because of this and its performance ratio of >1.0 as previously discussed, most internal spacecraft components are painted black. Most external spacecraft surfaces are painted
white to minimize solar energy absorptance.


The main disadvantage of coating is that it degrades quickly due to the operating environment.


Thermal control surfaces are also a key part of other thermal control methods which include multi-layer insulation (MLI) blankets and radiators. Aluminized Kapton is commonly used for the MLI blanket external layer. Surfaces with a performance ratio of less than 0.4, like white paints or optical solar reflectors (OSRs), make them effective radiators. OSRs include quartz mirrors, silvered or aluminized Teflon.



This is the primary system of heat rejection on spacecraft. Device with a large surface area used to radiate heat into space. Sizing depends on heat loads and temperature requirements.

A radiator can utilize an existing part of the spacecraft structure that is exposed to space for heat dissipation or it can be an individual component of the thermal control system that is designed to be a radiator. In both cases, internal excess heat is directed to the radiator for rejection into space as infrared (IR) radiation, thereby cooling the spacecraft.

Radiators come in several different forms, such as spacecraft structural panels, flat-plate radiators mounted to the side of the spacecraft, and panels deployed after the spacecraft is on orbit.

An important part of any radiator is its surface coating as previously described. This coating should have a high emissivity (e.g. white or OSRs) to increase heat dissipation efficiency. In addition, it should have a low absorptivity and be as stable as possible. Stability refers to its resistance to thermal degradation due to the space environment.

Most spacecraft radiators reject between 100 and 350 W of internally generated electronics waste heat per square meter. Radiators’ weight typically varies from almost nothing, if an existing structural panel is used as a radiator, to around 12 kg/m2 for a heavy deployable radiator and its support structure.


Multilayer insulation (MLI)

Thermal insulation is used as a thermal radiation barrier from incoming solar or IR flux and/or to prevent undesired radiative heat dissipation. Multilayer insulation (MLI) is the most common passive thermal control element used on spacecraft.


MLI prevents both heat losses to the environment and excessive heating from the environment. Commonly used to maintain temperature ranges for electronics and batteries in-orbit, or more recently, for biological payloads, thermal insulation is usually in the form of MLI blankets. However, the use of metallized tapes is also common for small spacecraft applications. Spacecraft components such as propellant tanks, propellant lines, batteries, and solid rocket motors are also covered in MLI blankets to maintain the ideal operating temperature.


MLI consist of an outer cover layer, interior layer, and an inner cover layer. The outer cover layer needs to be opaque to sunlight, generate a low amount of particulate contaminates, and be able to survive in the environment and temperature to which the spacecraft will be exposed. Common environmentally degrading effects on thermal control surfaces include UV radiation, atomic oxygen, charged particles (e.g. electrons), and spacecraft contamination (e.g. due to attitude control jets and/or propulsion out gassing).  Some common materials used for the outer layer are fiberglass woven cloth impregnated with PTFE Teflon, PVF reinforced with Nomex bonded with polyester adhesive, and FEP Teflon.


The general requirement for the interior layer is that it needs to have a low emittance. The most commonly used material for this layer is Mylar aluminized on one or both sides. The interior layers are usually thin compared to the outer layer to save weight and are perforated to aid in venting trapped air during launch. The inner cover faces the spacecraft hardware and is used to protect the thin interior layers. Inner covers are often not aluminized in order to prevent electrical shorts. Some materials used for the inner covers are Dacron and Nomex netting. Mylar is not used because of flammability concerns. MLI blankets are an important element of the thermal control system.


MLI is delicate and performance drops drastically if compressed (causing a thermally conductive “short circuit”), so it should be used with caution or avoided altogether on the exterior of small satellites that fit into a deployer (e.g., P-POD, NLAS). MLI blankets can also pose a potential snagging hazard in these tight-fitting, pusher-spring-style deployers. Additionally, MLI blankets tend to drop efficiency as their size decreases and the specific way they are attached has a large impact on their performance.


Due to this, MLI generally does not perform as well for small spacecraft (more specifically CubeSat form factors) as on larger spacecraft. Surface coatings are typically less delicate and are more appropriate for the exterior of a small spacecraft that will be deployed from a dispenser. Lastly, internal MLI blankets that do not receive direct solar thermal radiation can often be replaced by a variety of low emissivity tapes or coatings that perform equally well in that context, using less volume and at a potentially lower cost. Second-surface silvered FEP tapes offer excellent performance as radiator coatings, reflecting incident solar energy while simultaneously emitting spacecraft thermal energy efficiently, but the tapes must be handled carefully to maintain optical properties and they don’t always bond well to curved surfaces.


Active thermal control system (ATCS)

Ideally, you want all thermal control devices to be passive in order to maximize electrical power to the payload. However, this is not practical due to the extreme cold/hot environments “felt” by the spacecraft. Also, for some high-power spacecraft components, passive elements alone cannot continuously maintain their operating temperature range. This is where active thermal control is typically employed, to cover those situations where passive thermal control is inadequate.


Many active control methods also require continuous monitoring of spacecraft temperatures. To help achieve this, most spacecraft are covered from top to bottom with temperature measuring devices called thermistors. A thermistor is an electrical component that changes its resistance as its temperature changes. Since thermistor resistance is a function of temperature, the local temperature can be measured where the thermistor is located. These temperature measurements can then be used as inputs to an active control system and/or used for thermal trend analysis by an engineer.


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.


Heat Pipes

Heat pipes use a liquid-gas phase change to efficiently transfer heat from one location to another over relatively long distances. Most consist of the following components: evaporator, tube, working fluid, capillary wick structure, and a condenser. Heat pipes function by adding heat to one end of the tube at the evaporator, changing the working fluid from a liquid to a gas, and then removing the heat transported to the other end by the condenser, changing the working fluid back into a liquid.


Using this closed two phase system for thermal control, large amounts of heat can be transferred by convection through the sealed heat pipe tube between two end locations. The operating temperature range to be “felt” by the pipe determines the working fluid to be used which must
also be compatible with the material used for the tube. An aluminum tube with an axial groove wick structure, using ammonia as the working fluid has been commonly used for heat pipes.


Pumped Fluid Loops

Like heat pipes, pumped fluid loops (PFLs) can also be used to transfer large amounts of heat, to provide heat transfer cooling by forced convection. The cooling concept is similar to how the coolant is used in your car to cool the engine, where antifreeze is used as the coolant, cooling the engine as it passes through it then releasing heat as it flows through the radiator. The working fluid (e.g. ammonia, water) in the tubing absorbs thermal energy at a heat source and transfers it by mechanical pump to a heat sink.



Louvers are active thermal control elements that are used in many different forms. Most commonly they are placed over external radiators, louvers can also be used to control heat transfer between internal spacecraft surfaces or be placed on openings on the spacecraft walls. A louver in its fully open state can reject six times as much heat as it does in its fully closed state, with no power required to operate it. The most commonly used louver is the bimetallic, spring-actuated, rectangular blade louver also known as venetian-blind louver. Louver radiator assemblies consist of five main elements: baseplate, blades, actuators, sensing elements, and structural elements.


Offer a controlled rate of heat transfer, but can result in high temperatures if pointed toward sun. Second Surface Mirrors are more cutting edge and have all but replaced louvers in industry. Instead of louvers with different coatings, mirrors act to reflect incident radiation while radiating out internal energy



Heaters are used in thermal control design to protect components under cold-case environmental conditions or to make up for heat that is not dissipated. Heaters are used with thermostats or solid-state controllers to provide exact temperature control of a particular component. Another common use for heaters is to warm up components to their minimal operating temperatures before the components are turned on.

The most common type of heater used on spacecraft is the patch heater, which consists of an electrical-resistance element sandwiched between two sheets of flexible electrically insulating material, such as Kapton. The patch heater may contain either a single circuit or multiple circuits, depending on whether or not redundancy is required within it.

Another type of heater, the cartridge heater, is often used to heat blocks of material or high-temperature components such as propellants. This heater consists of a coiled resistor enclosed in a cylindrical metallic case. Typically a hole is drilled in the component to be heated, and the cartridge is potted into the hole. Cartridge heaters are usually a quarter-inch or less in diameter and up to a few inches long.

Another type of heater used on spacecraft is the radioisotope heater units also known as RHUs. RHUs are used for travelling to outer planets past Jupiter due to very low solar radiance, which greatly reduces the power generated from solar panels. These heaters do not require any electrical power from the spacecraft and provide direct heat where it is needed. At the center of each RHU is a radioactive material, which decays to provide heat. The most commonly used material is plutonium dioxide. A single RHU weighs just 42 grams and can fit in a cylindrical enclosure 26 mm in diameter and 32 mm long. Each unit also generates 1 W of heat at encapsulation, however the heat generation rate decreases with time. A total of 117 RHUs were used on the Cassini mission.


Design Process

  • establish the thermal requirements
  • establish the worst case for environmental heat loads and power dissipation
  • elaborate the control means: Begin with Passive System, adding components as needed
    Use active system if there is only a few degrees of tolerance in the required temperature or if several kilowatts are to be dissipated
    Design is typically modeled for the coldest case
  • build mathematical models to simulate the satellite thermal behaviour
  • analyze the design for worst environmental and dissipation heat loads
  • verify the design against the requirements
  • gather the budgets
  • change the design if necessary
  • verify the design by test and correlate the mathematical models.

Quality assurance

ESA has developed STEP-NRF (Network-model Results Format) and STEP-TAS (Thermal Analysis for Space) as open standards for product data exchange based on the ISO 10303 (better known by its informal name STEP, Standard for the Exchange of Product model data).
The thermal control project must be compliant with international standards of quality assurance from the methodology applied, to the analysis performed, and the test used for validation. Some relevant international standards from the European Cooperation for Space Standardization (ECSS, are:
• ECSS-E-30 Part 1. Space engineering. Mechanical — Part 1: Thermal control
• ECSS-E-10-03A. Space engineering. Testing
• ECSS-E-10-04A. Space engineering. Space environment



Future of thermal control systems

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
Advanced spray on thin films
Silvered quartz mirrors
Advanced metallized polymer-based films





References and Resources also include:


Cite This Article

International Defense Security & Technology (February 2, 2023) Spacecraft thermal control subsystems (TCS) essential for both the physical integrity of the satellite and for its efficient operation. Retrieved from
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"Spacecraft thermal control subsystems (TCS) essential for both the physical integrity of the satellite and for its efficient operation." International Defense Security & Technology [Online]. Available: [Accessed: February 2, 2023]

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