The first part of this article reviews the design drivers and the technologies currently used for spacecraft thermal control. The second part focussing on future technology developments in thermal control will appear in a later issue of the Bulletin.
Why the need for thermal control?
The need for a Thermal Control System (TCS) is dictated by the technological/functional limitations and reliability requirements of all equipment used onboard a spacecraft and, in the case of manned missions, by the need to provide the crew with a suitable living/working environment. Almost all sophisticated equipment has specified temperature ranges in which it will function correctly. The role of the TCS is therefore to maintain the tempera-ture and temperature stability of every item onboard the spacecraft within those pre-defined limits during all mission phases and thereby using a minimum of spacecraft resources.
The overall thermal-control function can be split into several different sub-functions (Fig. 1).
Figure 1. Interactions between the TCS sub-functions
Interaction with the environment
The external surfaces of a spacecraft may either need protection from the local environment or improved interaction with it, involving:
Heat provision and storage
In some cases, to reach or maintain the desired temperature level, heat has to be provided and/or a suitable heat-storage capability has to be foreseen.
In many cases, dissipated heat has to be removed from the equipment in which it is generated to avoid an undesirable increase in the unit's, and/or the spacecraft's temperature.
Generally speaking, it is not possible to reject the heat directly where it is generated, and appropriate means have to be used to transport it from the collection device to the radiating device.
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.
The design drivers
The major parameters driving the TCS design are:
About the environment
For all spacecraft, the incoming energy from the Sun and the heat radiated to deep space are usually the major environmental interactions. Depending, however, on the spacecraft's orbit and attitude, other para-meters can have an important influence on the thermal- control design. For example, the type of attitude stabilisation used can affect the TCS design. In general, spin stabilisation is more benign, as the rotation induces an averaging of the environmental flux input. Three-axis stabilised spacecraft need greater protection against short-term variations in energy input from the Sun or Earth.
Low Earth Orbit (LEO)
This orbit is frequently used by spacecraft that monitor or measure the characteristics of the Earth and its surrounding environment (Earth observation, geodesy, etc.), and by unmanned and manned space laboratories (Eureca, the International Space Station, etc.). The orbit's proximity to the Earth has a great influence on the TCS 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 h) and long eclipse duration (up to a third of the time). Small instruments or spacecraft appendages such as solar panels that have low thermal inertias can be seriously affected by this continuously changing environment and may require very specific thermal design solutions.
Ascent and re-entry
For space transportation systems, the ascent to and re-entry from operational orbit (usually, LEO) can introduce additional TCS design constraints. During these two phases, the environment is often too warm to reject heat by radiation, and the radiators used in orbit are often closed or protected. Consequently, alternative heat sinks (e.g. flash evaporators) or special TCS designs providing a high thermal inertia have to be foreseen to manage these heat loads.
Geostationary orbit (GEO)
In this 24 h 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 h 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 (Exosat, IRAS, ISO, etc.), 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, etc.
Missions designed for the long-term observation of particular phenomena need a constant, stable environment and therefore tend to make use of stable orbits needing very low station-keeping resources, far away from any celestial body, e.g. around a Lagrangian point. Scientific spacecraft such as SOHO and the future scientific mission COBRAS- SAMBA, are typical of this class of mission. The spacecraft are Sun-pointing and therefore have one side constantly illuminated and all other faces exposed to deep space. Consequently, the TCS design can be optimised fairly easily, unless there are very special temperature requirements or there is insufficient electrical power for the heaters.
In particular, spacecraft with cryogenic payloads benefit from a low and stable temperature environment in terms of mass (if cryostats are used) or power and complexity (for satellites using cryo-coolers).
Deep space and planetary exploration
This class of mission includes many different sub-scenarios depending on the particular celestial body or target exploration zone. 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/energy required for that survival phase.
About the heat dissipation and its
Two factors are important in this context for the TCS design, the absolute value of the heat to be dissipated and its distribution onboard the spacecraft, i.e. the power density. The first value has a big impact on the heat- rejection function (the dimensions of the radiator area increase with increasing power), while the power density defines the heat collection and transport functions (a high power density calls for highly efficient heat removal). Typical installed powers for various kinds of spacecraft are compared in Table 1.
Installed Power (W) Mission Orbit Attitude min. max. Science: - astronomy HEO, Fixed point Sun pointing (mostly) - deep space Various transfer orbits Sun or planet pointing 200 1 500 Telecommunication GEO Earth pointing 500 5 000 Earth Observation LEO Earth pointing 500 5 000 Meteorology GEO Earth pointing 200 1 500 Manned Vehicles Transfer +LEO Various 1000 10 000 Manned Stations LEO Sun pointing 3000 30 000
Two conflicting requirements can be detected in terms of power utilisation:
Another very important factor is the duty cycle. The best solution would be power dissipa-tion which compensates for the change in environmental fluxes (e.g. maximum power dissipation during eclipses!), so as to have an almost constant global heat input to the spacecraft. Given the present, near- and probably medium-term power generation methods, the reality is exactly the opposite: the maximum power dissipation occurs together with the maximum environmental fluxes. This forces the TCS design towards an over-dimensioning of the heat transport and rejection equipment to cope with the concurrent peaks. In turn, this over- dimensioning causes an increase in the complexity of the design and the need for more resources during the cold phases of the mission.
This introduces the third mutual interaction between the power subsystem and the TCS, namely the availability of power during cold mission phases for the heat-provision function. During those phases, power is generally provided by batteries and is therefore limited. This limitation can create further complexity in the TCS design.
About the temperature requirements
This factor is largely related to the technology of the spacecraft equipment. As already mentioned, it is the TCS's task to keep all equipment items working within their allowable temperature ranges, which in turn depend on the internal design, the components used and, last but not least, the required reliability. This applies in particular to electronic and electro-mechanical equipment, the design of which is often too similar to that of its 'terrestrial' counterpart, which has to work in a much more benign environment (air is an added value for TCS!). Improved thermal designs, in combination with better definition of the allowable temperature ranges, could save projects both time and money in the long run.
Three relevant temperature bands can be defined:
We will concentrate here on the 'conventional range', articles relating to the other two ranges already having been published in past editions of the ESA Bulletin (e.g. No. 75, August 1993 and No. 80, November 1994).
Within our conventional range, different sub-ranges can be identified according to the various equipment requirements. Classical examples include:
Nonelectronic items can have a wide range of temperature requirements, most of which are of a functional nature, e.g. limitation of thermal noise in sensors. Some extreme examples are shown in Table 2.
Operating Non-oper./Storage Uniformity Stability Item Temperatures (°C) Temperatures (°C) (°C/m) (°C/min) Min. Max. Min . Max. Video-Camera CCD -150 -100 - - - ±0.5 Laser Thermal I/F 5 10 5 10 ±0.5 ±0.1 Fluid-Physics Samples 5 90 5 40 ±0.1 ±0.01 Life-Science Samples 4 38 -80 -80 ±1.0
Temperature uniformity and stability can have an even greater impact on the TCS design than the absolute temperature values themselves. The former can be expressed either as the maximum allowable temperature difference between two adjacent parts, or as the maximum temperature gradient in continuous bodies. The temperature stability refers to the maximum allowable variation in the temperature of a specific item over time. The ability to cope with these requirements depends on the environment and power-duty-cycle design drivers and on the actual spacecraft configuration.
Great care has to be taken to discriminate between the 'nice to have' and the truly mandatory requirements, as sometimes even few degrees (or a few tenths for stability) can make the difference between a feasible and an unfeasible system or, at the very least, between an affordable and a very expensive system.
About the spacecraft configuration, reliability and
One of the major problems of TCS design is that the spacecraft configuration is usually defined based on the physical accommodation of the various payload and basic subsystem (propulsion, solar arrays, etc.) elements. Only when the physical configuration is virtually frozen is the TCS designer brought in to assess whether all of the temperature requirements can be met. Should this not be the case, a great deal of time (and money) then has to be spent in trying to re-accommodate equipment and find ad-hoc solutions, which are never resource-efficient. Concurrent engineering should be applied more often at all levels, from the equipment to the spacecraft design, to try to overcome these not infrequent problems.
Reliability affects the TCS both directly (the TCS function has its own requirement) and indirectly through the equipment temperature requirements. The greatest impact is on the heat- provision, transportation and rejection functions. For manned vehicles, for instance, the reliability required for the cooling loops can cause an enormous increase in the complexity and mass of the TCS.
Verification, and in particular testing, require-ments have too often been the cause of an efficient TCS design being rejected. The reluctance to use heat pipes due to the complications introduced into the thermal system tests (see section on heat-transport systems) is a classical example. As already demonstrated by many commercial spacecraft, a proper combination of testing at component and system level with analytical correlation techniques can solve such problems, resulting in a simpler and more efficient temperature control system.
Importance of parameters
The various design drivers have different impacts on the various TCS functions and on the mass, complexity and cost of their respective design solutions. Table 3 gives an indication of the relationships between the design drivers examined and each TCS function ('o' means little or no impact, while an 'x' indicates a growing level of importance; M = mass; CX = complexity; CT = cost). Environment Protection Heat Prov. and Storage Heat Collection Heat Transport Heat Rejection
Environment Heat Prov. Heat Heat Heat Protection and Storage Collection Transport Rejection Design Drivers M CX CT M CX CT M CX CT M CX CT M CX CT Environment xx xx xx x xx x o o o o o o xxx xx xx Heat Dissipation - absolute o o o o o o xx x xx xx x xx xxx xxx xxx - density o o o o o o xx xxx xxx xx x xx x x x Temperature - level x x x xx xx x x x x x xx xx xxx x xx - stability x x x xx xxx xx xx xx xx xx xx xx x x x - uniformity x x x xx xxx xx xx xx xx xx xx xx x x x Reliability o o o x xxx xx xx xx xx xx xx xx xx xx xx Configuration x x x x x x x xxx x xx xx xx xx xx xx Assembly, x x x xx x x xx x x xx xx xx xx x x Integration
Interaction with the
The easiest way to modify a surface's behaviour is to coat it with paint or a layer of other suitable material. All spacecraft make use of many different kinds of coatings, ranging from relatively simple- to-apply paints to more sophisticated chemically or physically produced conversion coatings. Coatings are characterised by their thermo-optical properties: absorptivity, emissivity, reflectivity and transparency.
The main disadvantages of coatings are the degradation caused by the operating environment and the contamination induced by ground handling or space operations, the absorptivity being the parameter most affected. Both the handling on the ground and the space environment tend to increase the initial absorptivity of a coating towards an end-of-life (EOL) value. The latter depends on the time spent in orbit, the relevant environment (particle fluxes, UV flux, etc.) and the orientation of the surface vis-a- vis the motion of the spacecraft.
A correct TCS design has to duly take into account all of these factors and use suitable begin-of-life (BOL) and EOL values.
Multi-Layer Insulation (MLI)
When a simple coating is insufficient to avoid great heat losses or gains for a surface, multi-layer insulation can be used. It consists of a certain number of layers of plastic material (normally Mylar or Kapton) coated on one or both sides with a layer of metallic material to reduce the radiation, and separated by sheets of spacer material (e.g. Dacron net) to avoid direct contact between adjacent foils. The external foil coating depends on the particular application: it can be painted or metallised, or can even consist of a different material (e.g. glass-reinforced cloth).
MLI efficiency can be defined either in terms of the linear conductance through the blanket, or via the so-called 'effective emittance'. In the first case, the thermal flux can be calculated as the product of the given value times the temperature difference between the external layer and the hardware covered by the blanket. In the second case, it is calculated as a radiative heat exchange using the effective emittance (Fig. 2). This parameter has a very simple mathematical formulation, but it can have quite different physical meanings and the choice of definition depends on the modelling technique used.
Figure 2. Effective-emittance definition for various MLI layouts
The factors affecting the efficiency are the physical composition of the blanket (number of layers, type of coatings, etc.), the average blanket temperature (usually the arithmetic mean between the two outermost layers), the eventual presence of air or humidity within the layers, and the pressure between them. A very important factor is the way in which the blanket is applied to the spacecraft surface: a single piece of blanket covering a large surface is more efficient than a number of small blankets covering the same surface. A blanket suspended over a surface (case 3 of Fig. 2) is more efficient than one in direct contact with the surface (case 1 of Fig. 2).
Generally speaking, the MLI's efficiency is measured on relatively small samples, while the real efficiency of an MLI system is only known at the time of system-level thermal testing. Consequently, suitable safety factors have to be applied during the design phase.
Figure 3 shows the thermal conductance versus temperature for MLI samples measured at ESTEC for some recent ESA programmes. Figure 4 shows the thermal conductance versus the average temperature for samples and the real (with overlapping, seams, etc.) MLI (having the identical composition) as measured for Spacelab.
Figure 3. Thermal conductances of several MLI samples as a function of average temperature
Figure 4. Effect of overlapping and presence of gaps on MLI thermal conductance
A surface may only need to be protected during particular mission phases, whilst at other times it has to be free to radiate to deep space. Louvres can be used either to provide a heat sink during phases with Sun illumination, or to reduce the heat losses during cold (shadow) phases.
In the louvred radiator shown in Figure 5a, each blade is provided with a sensor/actuator element (e.g. a bi-metallic spring) which senses the temperature of the radiator baseplate and rotates the blade accordingly. The radiator can be blocked off completely when the temperature is lower (or higher for Sun louvres) than a pre-defined value, and exposed to various extents depending on the prevailing temperature levels. The accuracy of the temperature regulation depends on the physical characteristics of the louvre mechanism and is generally limited to ±5°C.
Figure 5. Louvre (a) and shutter (b) schematics
Louvres for use on top of radiators were developed in Europe in the early 1970s by ERNO and SNIAS (today DASA Aerospace and Aerospatiale, respectively), but they have not often been used aboard European spacecraft.
A shutter (Fig. 5 b) consists of a thin metallic plate (or blanket), which can be slid over a surface (usually by an electric motor) to vary the exposed radiator area in an almost continuous manner from zero to maximum exposure. The advantages compared with louvres are a greater effective emissivity when the shutter is completely open (no or very limited multi-reflection effects) and a better insulation efficiency when completely closed. A thermal shutter of this type was used on ESA's Giotto spacecraft.
The advantages of louvres and shutters are greater adaptation to the environmental conditions and a reduction in the power and energy needed for heating during cold phases. The disadvantages are the mass and presence of the associated mechanisms, which can reduce the reliability of the TCS.
Electrical resistance heaters are the easiest means of providing heat to spacecraft equipment. The provision and storage functions are separated in that the former is performed by the TCS, while the latter is provided via the power subsystem.
The heaters can be powered continuously or, more usually, can be switched on and off according to the temperature of the controlled element. In the latter case, it is possible to have local control by thermostats or central control via a dedicated switching unit (the so-called thermal-control unit) or via the spacecraft Data Handling System (DHS). This implies the use of temperature sensors and data and command lines. Depending on the particular spacecraft configuration and temperature requirements, this heater monitoring and commanding system can become quite complex. The main disadvantages of heaters are therefore the need for electrical energy and either the complexity of the DHS or the reduced reliability if thermostats are used.
Electrical heaters are used on all spacecraft. In recent years, European heaters have been qualified according to the very stringent ESA specification for both single- and double-density designs (up to 200 Ohm/cm²).
Some planetary and exploratory missions to the periphery of the Solar System cannot rely on the Sun and batteries to produce and store electrical power for TCS purposes. Radio-isotope Heating Units (RHUs), based on plutonium, have then been used either to warm the spacecraft directly or to produce electricity via Radio-isotope Thermoelectric Generators (RTGs) to feed the heaters. There are currently no European manufacturers of RHUs or RTGs, but both the USA and Russia have developed and used these devices for their deep-space missions. Political as well as procurement problems will make the use of this type of RHU less and less acceptable in the future.
Phase Change Materials (PCM) offer the possibility to store thermal energy directly as latent heat of fusion or sublimation. The item to be controlled is linked to a vessel filled with a PCM. When the item is active, the PCM absorbs heat and melts or sublimates at a stable temperature; when the equipment is inactive the PCM can resolidify, releasing the relevant amount of heat. Usually, the melting PCMs can be used easily in reversible, closed systems, while the sublimating PCMs are used in open, non-reversible systems (i.e. the gas is released after the phase change to avoid over-pressure).
The most critical parameters are the temperature at which the phase change takes place, and the amount of heat absorbed or released during the change. The temperature ranges normally of interest are the near-zero range ( 5 to +10°C), or particular ranges for specific experiments, e.g. 80°C for life-science experiments. Other important parameters are the thermal conductivity and the density of the two phases; the former because of the need to transfer the heat efficiently inside the PCM, and the latter because the containing structures have to withstand the volumetric change of the PCM.
Two advantages of a PCM device are the stability of temperature control and the absence of moving parts. The heat- storage requirement is defined by the duty cycle for reversible systems, and by the total operating time for non-reversible (e.g. sublimating, boil-off PCMs) systems. Since the mass of the device is directly proportional to the heat-storage capability, it is difficult to use a PCM device without incurring a serious impact on the overall mass budget. Moreover, problems associated with the limited thermal conductivity of many PCMs make it necessary to use finned containers, which again increase the mass and volume of the devices. Another point of concern is the design of the container against leakage, for both safety (PCMs can be quite corrosive) and functional reasons.
PCM-based devices have been used on US spacecraft, including some Shuttle-launched missions. Various breadboards were developed in Europe in the 1970s but, apart from an application on Spacelab, there are no references to their use aboard other European spacecraft.
Heat collection and transportation
The selection of the most appropriate system and components depends on the overall power level, the power density and the temperature requirements.
The usual way to collect the heat dissipated by any item of equipment is through its baseplate and fixation elements (mounting feet). With increasing power dissipation, the complete baseplate has to be in contact with the spacecraft panel. The heat transferred then depends on such parameters as the interface pressure, the surface finish, the types of materials involved, etc., which are sometimes difficult to quantify (at design level) and control (during manufacture and integration). Ways of increasing the conductivity through the interface surfaces include the use of metallic or synthetic mats, or the application of thermal grease. This last solution has to be used with care because of the obvious potential contamination problems.
In some cases, a number of units are connected together to an intermediate solid panel called a 'doubler', which is usually made of aluminium. This doubler spreads the heat dissipation over a larger area, thereby providing an improvement in the temperature uniformity and an increase in the effective contact area to the heat-transport or heat-rejection device. It is convenient to locate redundant units or units operating with different duty cycles on the same doubler in order to use the heat dissipated by the operating units to maintain the others within limits without the need for additional heating power. The disadvantage of this simple solution is the mass of the doubler, which must be reasonably thick to achieve a good efficiency.
Braids of conductive material (e.g. copper) are sometimes used to connect heat-dissipating equipment to a 'remote' radiator. As the overall conductance is proportional to the braid's cross- section and inversely proportional to its length, this method can clearly only be used for short distances and very low heat loads. As an example, a copper rod weighing about 22 kg would be needed to transport 10 W over a distance of 1 m with a temperature difference of 10°. For comparison, a simple heat pipe (e.g. a stainless-steel/ammonia heat pipe with a diameter of 9.5 mm) provides a better performance (lower temperature difference) for a mass of 0.25 kg/m, i.e. about 100 times less. One advantage of the braid is its flexibility, which provides a certain degree of isolation from vibration and helps to avoid configuration problems.
The heat pipe is a device that allows efficient transport of thermal energy. It typically consists of a sealed metal tube with a capillary structure on the inside, filled with a suitable working fluid. Heat is absorbed at one end by evaporation of the fluid, and released at the other by condensation of the vapour. The liquid is transported back to the evaporator by capillary forces.
The heat pipes most commonly used on spacecraft are an aluminium/ammonia type that allows optimal temperature control in the 0-40°C range. Since the quantity of heat transported by the pipe is defined by its design and dimensions, the equivalent thermal conductance is fixed, leading to the Constant Conductance Heat Pipe (CCHP in Fig. 6a).
Figure 6. CCHP (a) and VCHP (b) schematics
There is also a special type of heat pipe known as a Variable Conductance Heat Pipe (VCHP, Fig. 6b). This device provides better temperature control when the equipment can either dissipate at different power levels, or the condenser is exposed to a varying environment. The amount of heat transferred is usually controlled by blocking part of the condenser area with an inert gas.
As the capillary forces are weaker than gravitational forces, heat pipes can only work in a gravity field if the evaporator and condenser are at the same level, or if the evaporator is below the condenser (so-called 'reflux mode'). Consequently, if a spacecraft has heat pipes located in different planes, it is not always possible to fully verify the complete thermal design with system-level testing alone. However, as already mentioned, this constraint can be overcome and should therefore not restrict the use of heat pipes, which offer great advantages.
For greater power dissipation or more stringent temperature requirements, other heat collection and transportation systems can be used. Various kinds of fluid loops have been proposed and applied to cope with these situations.
In single-phase loops, the refrigerating liquid absorbs the heat from the heat-dissipating items (e.g. via a cold plate or a heat exchanger) by increasing its temperature, and transports it to the heat-rejecting device (heat exchanger or directly via a radiator) where the fluid is cooled down. A mechanical pump is essential to provide the hydraulic energy needed for this task (Fig. 7a).
Figure 7. Schematics of cooling loops: (a) Single-phase loop. (b) Mechanically assisted two-phase loop. (c) Two-phase capillary loop. (d) Two-phase hybrid loop
The advantages of these systems lie in their flexibility and lack of sensitivity to their orientation and mechanical environment. The fluid flow rate can be easily regulated (e.g. via a variable-speed pump), allowing either a wide range of power duty cycles (a ratio of 1 to 10 is possible) and/or different levels of temperature accuracy, stability and uniformity. The temperature range can be adapted to the particular application by selecting an appropriate fluid. Since the fluid is circulated by the mechanical action of the pump, the system operates with the same efficiency on the ground, onboard a spacecraft, or during a descent on a celestial body. Disadvantages are the power needed to drive the pump and possible vibrations induced by the pump and fluid flows.
Single-phase fluid loops have been widely used since the early days of manned spaceflight. In Russia, they have also been used frequently for unmanned spacecraft; e.g. air loops have been used on Proton, liquid loops on high-power telecommunications spacecraft (in conjunction with deployable radiators), and combined liquid/air loops on retrievable low-orbit spacecraft (e.g. Foton). In Europe, they have been used on Spacelab and Eureca, and will be used in the future in the Columbus Orbital Facility as well as the Mini-Pressurised Logistic Module.
Two-phase Mechanically Pumped Loops (MPLs, Fig. 7b) are similar to single-phase loops, except that the fluid changes state (evaporating while absorbing the heat and condensing in the heat-rejecting devices) instead of just changing temperature. The advantage compared with the single-phase type lies in the much lower fluid flow rate needed to manage the same quantity of heat (due to the use of the latent heat of evaporation) and the associated reduction in the level of resources needed for the TCS (lower electrical consumption of the pump, lower mass due to smaller fluid lines and fluid inventory, etc.).
In Capillary Pumped Loops (CPLs: Fig. 7c), the driving force is provided by the capillary action of the wick material inside the evaporators and a separate mechanical pump is not needed. However, there are particular operations or mission phases for which assistance to the capillary action might be desirable (e.g. loop start-up, peak power loads, high mechanical loads or ground testing).
Hybrid loops (Fig. 7d) consisting of a CPL with a mechanical pump are now being proposed. During nominal operations, the pump is bypassed and the fluid flow is ensured by the capillary actions. Only during critical phases is the pump inserted into the loop to provide the extra energy needed by the fluid. Many experimental CPLs have been or are being flown to demonstrate the technology, which is currently baselined for several Earth- observation experiments, e.g. the European ATLID and the American EOS-AM.
These are used to transfer heat from a fixed spacecraft element to any deployable/ movable/rotating element (e.g. a radiator). Depending on the nature and extent of the allowable motion (single deployment, continuous rotation, etc.), the joint can either be very simple (the braid mentioned above for low heat loads) or considerably more complicated.
Flexible heat pipes have been proposed for single deployments, and rotating thermal joints (based on shape-memory alloys or gas pressure) for periodic rotation. They have yet to be flown on a European spacecraft.
A radiator is simply a (highly) conductive panel exposed to deep space and (normally) coated with a high-emissivity coating. Depending on the spacecraft's size and configuration, there can be central radiators to which all the heat dissipated on board is transferred, or multiple radiators each dedicated to a payload unit or group of payloads and/or subsystems.
The dissipating equipment can either be mounted directly on the radiator or connected to it via heat pipes or fluid loops. In the latter case, the heat pipes or fluid lines can either be fixed to the external faces of the radiator or directly embedded into its structure. The second solution is more efficient from the structural (mass-saving) and thermal viewpoints, but can also be less reliable due to the probability of micro-meteoroids impacting the radiator, and is more critical with regard to spacecraft integration activities.
The radiator's size depends on the power to be dissipated, the temperature of rejection (defined by the items to be controlled) and the temperature of the surrounding environment (Fig. 8). In most cases, the radiator is mounted on a spacecraft panel and therefore only radiates on one side. In the case of high and/or varying powers or changing environmental conditions, this configuration is not very efficient. A better solution is to use both faces of the radiator, but this implies the need for radiator deployment.
Figure 8. Influence on radiator area of environmental (sink) and radiator temperatures
One way of coping with changing heat loads is to use louvres or shutters on top of the radiator, as discussed earlier.
Thermo-electric heat pumps
Heat pumps are reversible machines able to transfer thermal energy from lower temperature to higher temperature bodies using an additional source of energy. Only thermo-electric heat pumps have been used in space until now, the basic feature of which is the Peltier element, which results from the connection through a metallic tab of type-n and type-p semiconductor materials.
The efficiency of a Peltier element depends on its intrinsic characteristics (thermo-electric effect, thermal and electrical conductivity), the electrical current, the temperature to be controlled and the temperature of the heat sink. The overall performance of a thermo-electric heat pump is strictly related to the efficiency of the thermal coupling between the Peltier element tabs and the surfaces to be cooled or heated.
For low cooling/heating loads, the elements are bolted in between the baseplate of the controlled element and the heat sink. Thermal grease is usually applied at the interface to increase the joint's thermal efficiency. However, as the interface pressure cannot be high for mechanical reasons, this method is not adequate when high thermal performance is required (very strict temperature control and/or high cooling/heating loads). The preferred solution in this case is soldering of the elements to the heat sink.
The most efficient heat sinks are currently water heat exchangers. Good performance can also be obtained with air heat exchangers, at the expense of larger volumes and higher power consumptions (needed to drive the fans). In all other cases, the cooling/heating loads as well as the temperature difference between the cold and hot side have to be very small, otherwise the electrical power required becomes prohibitive.
Thermo-electric heat pumps are normally used for tight temperature control of low-power instruments (advantages being the absence of vibration and ease of installation) and of facilities used for microgravity experiments. Many systems have been developed and used for both manned (e.g. ESA's Biorack) and unmanned spacecraft (e.g. Biobox aboard Foton).