Energy supply system (satellite)
As energy supply system or on-board power supply of a satellite are all systems for the generation, conversion , store- and distribution of energy called on board satellites. In English it is called Electric power / distribution subsystem (EPS or EPDS).
overview
With energy mainly is electrical energy to supply the control and regulation systems, as well as the payload (for example, sensors, receive and transmit electronics), and other subsystems (for example, by electric drive systems, and life support systems at manned satellites) meant. The heating and cooling (e.g. from sensors and on-board electronics) is often fed by the electrical system, but is implemented as an independent temperature control system .
In most satellites, the energy supply is provided by solar cells with the support of secondary cells ( accumulators or fuel cells ) if there is sufficient brightness from the sun in the area close to the earth. For short periods of use or missions in which the use of solar cells is not possible (e.g. re-entry or landing emissions) as well as when starting, the energy can also come from batteries ( primary cells ) or fuel cells. For satellites that are further away from the sun and so the supply of radiant energy is too low or where solar cells cannot deliver the required power, nuclear power systems are also used, for example in the form of radioisotope generators that are considerably smaller than other designs .
The energy supply systems used are designed in such a way that they can cover the energy requirements of the satellite in all operating states (during start-up, in normal operation with the payload systems switched on and off, during orbit changes, ...). Fault conditions and malfunctions must also be taken into account through redundancy and appropriate design. As is usually the case in space travel, the energy supply systems are faced with a trade-off between energy requirements, reliability and costs (for example in terms of weight and size).
functionality
The energy supply of a satellite normally consists of four parts (subsystems). These are energy generation, energy conversion or processing, energy storage with the corresponding charging and discharging electronics, and energy distribution. These systems do not necessarily have to be present, since, for example, when using nuclear energy or pure primary cells as an energy source, an additional energy storage device is not necessary. The opposite case can also occur if the system has been designed redundantly for safety reasons (for example an emergency power supply with primary cells is additionally installed) or for some subsystems of the satellite (for example the apogee motor ) an independent power supply adapted to the subsystem (for example again in form of primary cells) is used.
Energy generation and conversion
A basic distinction is made between energy generation based on the primary energy source. These can be external energy sources such as solar radiation or the field energy of magnetic fields (sun or planets). But it can also be internal energy carriers such as chemical energy carriers and nuclear fuels . The advantage of external energy sources is that the corresponding fuels do not have to be carried. The disadvantage is usually the limited performance or availability.
The conversion of the available primary energy takes place either directly (as in the case of solar cells, primary cells or fuel cells, for example), but can also take place indirectly via the generation of heat, which, however, was previously only used in nuclear power plants. The resulting heat can be converted into electricity via turbines or MHD generators, as with conventional systems installed on earth , but closed circuits (similar to the primary circuit of nuclear power systems) must be used, since satellites do not of course have an arbitrarily large amount of coolant is available.
Chemical energy
The generation of electrical energy through the reaction of chemical compounds is commonly known in the form of batteries (more precisely primary cells ). Fuel cells are also used in satellite technology (for example in the space shuttle ). Primary cells are often used for reserve and emergency energy supply , but also for short-term missions ( e.g. Lunar Roving Vehicle ) or to supply the systems of the launch vehicle or the apogee motor. Silver oxide-zinc (previously also mercury oxide-zinc ), lithium thionyl chloride , lithium sulfur dioxide and lithium carbon monofluoride batteries are often used. Thermal batteries are also used in some cases . Secondary cells (accumulators) are more used for energy storage (see below).
Solar energy
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Solar energy is the most common source of energy in satellites. Practically all geostationary and most research satellites use solar panels for energy supply. These convert the incoming sunlight directly into electricity through the barrier photo effect . The first satellite to use this technology was Vanguard 1 , which was launched on March 17, 1958. The most commonly used types are monocrystalline single-layer silicon cells (also high-eta cells with a structured surface) and multilayer gallium arsenide cells , with the former primarily utilizing visible light, the latter also using light from the infrared and UV spectrum, thus achieving higher levels of efficiency . At the beginning of their lifetime, these are 12 to 14% for silicon cells and around 25% for the gallium arsenide types. However, the efficiency is not constant. It falls under the influence of the high-energy (particle) radiation emanating from the sun and particularly sharply with increasing temperature, which can lead to an operating temperature between 60 ° C and over 100 ° C due to solar radiation. In addition, the performance values deteriorate over time as a result of these influences (which is called degradation ) by an order of magnitude of up to 30% for silicon cells and 13% for gallium arsenide cells with a service life of 15 years. This effect must be taken into account for satellites that are approaching the sun. In addition, influences from micrometeorites and material fatigue due to temperature fluctuations (when entering / leaving shadow areas or changes in alignment) must be taken into account or avoided.
The particle flow from the sun (especially during solar storms ) causes another effect that must be taken into account when designing solar cell booms. The surface of the satellite is electrostatically charged , which leads to potential differences of several hundred volts between neighboring surfaces at a rate of several volts per second (for example the front and back of the solar cells). This can lead to electrical discharges and damage to the solar cells. A short circuit (secondary arc ) of solar cells through the material knocked out during the (primary) electrical discharges must be ensured by an appropriate construction .
The solar cells themselves are either mounted directly on the surface of the satellites (for example, in the case of spin-stabilized satellites or satellites with low energy requirements) or designed as deployable or unrollable arrays. Arrays are usually designed to be fully rotatable in the unfolding axis in order to be able to track the solar radiation. The structural weight of these is in the order of magnitude of 100 watts / kg, whereby the corresponding unfolding and tracking mechanisms , as well as the energy storage devices necessary for weight and reliability, must also be taken into account. The achievable power depends on the size and the efficiency of the solar cell arrays and can be calculated from these values and the solar constant of about 1.37 kW per m 2 . It is in the range of a few watts for smaller satellites, over 10 kW for large communication satellites and over 100 kW for the ISS.
In addition to solar cells, since the pioneering days of space travel, solar dynamic systems for high performance have been planned time and again. The functional principle is similar to that of earthly power plants, but instead of primary energy sources (coal, oil) the sunlight bundled with the help of mirrors for heating and evaporation of the working materials (for example xenon , mercury or rubidium would be possible) and subsequent energy generation through Heat engines ( turbines , generators , recuperators , coolers and radiators ) are used. The physical principle of the circular process such as the Stirling , Brayton or Rankine process is used. The use of MHD generators would also be conceivable. Despite the theoretical advantages for larger systems, solar dynamic systems have not yet been used to supply energy to satellites and spaceships.
Nuclear energy
Nuclear energy as a primary energy source is mainly used in satellites that are further away from the sun or where solar cells cannot deliver the required power or are not practical due to their dimensions and properties. The advantages are their high reliability, long service life and compact dimensions. Their disadvantages are the necessary shielding of ionizing radiation and, above all, the problem of acceptance (see also Kosmos 954 ) of nuclear power plants, which for safety reasons must be designed in such a way that they can withstand an explosion of the launch vehicle or a crash.
As with solar cells, the systems used are divided into systems with direct energy conversion (static systems) and indirect energy conversion (dynamic systems). As with solar cells, however, dynamic systems (whose working principle is similar to that of terrestrial nuclear power plants with circular processes) have not yet been used.
Today mainly static systems are used, such as radioisotope generators (RTGs) based on the Seebeck effect, which are considerably smaller, lighter and simpler than other designs . These have an efficiency of about 5 to 10%, a weight of 10 to about 100 kg and are used in the power range of up to 1 kW electrical power. Some of these systems can also be used to control the temperature of the probes. Examples of RTGs are the SNAP systems from Ulysses , Galileo or Voyager and some Pioneer space probes. The first spacecraft with a radioisotope generator was the Transit 4A satellite , which was launched on June 29, 1961.
In some cases, such as the American test satellite Snapshot and the Russian RORSAT satellites, real nuclear reactors were used instead of the radioisotope generators, which operated with thermoelectric (RORSAT) or thermoionic (TOPAZ) energy converters . These have an efficiency of up to 25%, an electrical output of up to 100 kW and a much more complex structure and larger dimensions. They are discussed again for deep space probes with an electric rocket engine .
Other energy supply systems
In principle, other types of energy supply for satellites and space probes are also possible. The magnetic field of planets can also be used to generate electricity. These systems, also known as electromagnetic tether , are based on induction in kilometers of electrical conductors. Such systems have already been tested on the Space Shuttle flights STS-46 and STS-75 . Even more exotic proposals are the supply of satellites from Earth by cable (see space elevator ) or laser beam , as well as the use of photosynthesis (plant cultivation) in huge space stations.
Energy storage
With the exception of probes for special tasks, such as deep-space probes in which nuclear energy continuously supplies electricity, energy storage devices are essential for satellites equipped with solar cells in order to be able to maintain the energy supply to the satellite when there is insufficient solar radiation. This can be the case due to the shadowing of the satellite (for example by the earth, which is regularly the case with normal near-earth flight paths) or by incorrectly aligning the solar cells in the direction of the sun (for example in the event of control errors). Even when the satellite is started, in which the solar cells are usually transported in folded form for reasons of space, primary batteries or energy storage devices have to take over the supply until the solar cells unfold.
Secondary cells (accumulators) are mostly used as energy stores. Fuel cells (for example in the space shuttle) or flywheels are also used less frequently , although the advantage may be that the former can also be supplied from the spacecraft's fuel supply and the latter can also be used to stabilize the satellite. It should be noted that the energy storage devices used on orbits close to the earth must withstand up to tens of thousands of charging and discharging processes over the course of the lifetime of a satellite (sometimes several years). Furthermore, their parameters (such as capacity, voltage, permissible charging current or charging curve, internal resistance, ...) must be designed in such a way that they can meet the requirements of the satellite systems even at the end of their service life and always have a sufficient charge state through charging electronics and energy source have.
The secondary cells used are mainly nickel-cadmium , nickel-hydrogen and lithium-ion batteries , which became more common from the 2000s onwards . The service life of the accumulators depends not only on the period of use (calendar service life or storage service life), but above all on the number of charging cycles, the depth of discharge and the discharge current. Depending on these values, the nominal capacity and the nominal voltage decrease over time, while the internal resistance of the cells increases. However, operation outside of the specified application parameters (temperature, max. Discharge current, deep discharge or overcharging ) can greatly shorten the service life of the accumulators or even lead to their destruction (e.g. risk of explosion with lithium-ion cells in the event of overcharging) . -circuits must be prevented. Since the accumulators consist of several cells in series and / or parallel connection, the chargers and charging processes must be designed in such a way that the parameters are adhered to for all cells (see balancer ). The nominal voltage of the accumulators ranges from 1.25 to about 300 volts and capacities in the range from milliampere hours to more than 400 Ah (for example with the Hubble space telescope ).
Power distribution
The energy distribution system is an electronic component that provides and distributes energy (voltage supply) between primary and secondary energy suppliers (solar cells, RTGs, energy storage devices, ...) and the energy consumers (payload, charger for the energy storage device, satellite bus with thermal control, control systems,. .., but also load banks for excess energy). It also takes on monitoring, regulation and security tasks, so that the individual consumers can be supplied or switched off, both in normal operation and in the event of a fault, depending on the amount of energy available and the current operating status. Accordingly, the system must be designed to be flexible, fault-tolerant and robust (also with regard to radioactive radiation).
The power distribution system usually provides several voltage levels (for example ± 5, ± 12 and +28 volts) for the individual consumers, which can be stabilized and smoothed depending on the requirements . A distinction is made between regulated (BR), unregulated (BNR), semi-regulated (BSR) and hybrid (BH) on-board networks. An unregulated on-board network means supplying the on-board network directly from the built-in accumulators, the state of charge of this determining the voltage in the on-board network and the consumers having to cope with a correspondingly fluctuating supply voltage. The input voltage of the solar cells (and thus the charging voltage or current of the accumulators) is (contrary to the name) with this method limited by appropriate regulators. The advantage of this method is the simple structure of the system and the suitability for fluctuating and impulse-like loads, which is with the fluctuating supply voltage and the risk of a permanent breakdown of the on-board network (power lockup) due to operation of the solar cells under unfavorable operating conditions (practically short circuit of the solar cells by battery charge level is too low). With regulated on-board networks, the on-board voltage is obtained from the battery and solar cells via appropriate regulators. With a semi-regulated on-board network, this regulation only takes place if the solar cells are supplying sufficient electricity. With a hybrid bus, both regulated and unregulated on-board voltages (planes, sections) are available. In all processes, the regulation itself is carried out using an inverter , converter , load bank or voltage regulator . In solar cells as an energy supplier is still distinguishes between direct or indirect transfer of energy, with the first setting solar cells directly feed their power into the power supply bus and the latter via a DC-DC converter (DC-DC converter) to the maximum power point tracking principle.
The output voltage level for the systems is depending on the power 28 V (up to 3.5 kW), 50 or 65 V (up to around 10 kW) and around 100 V (if more than 10 kW, for example with the ISS) for regulated On-board networks and 28 V (up to 2 kW) and 35 or 42 V (over 2 kW) established for unregulated on-board networks.
Examples
| Surname | Type | Supply system | power | Mission duration | comment |
|---|---|---|---|---|---|
| VEGA | Launcher | Batteries | 1st stage: 48 Ah 2nd stage: 24 Ah 3rd stage: 8 Ah |
Minutes | |
| Space shuttle | Launcher | 3 fuel cells 3 nickel-cadmium batteries. 3 APU |
3 × 12 kW 3 × 10 Ah 3 × 5 kW (short-term max. 100 kW) |
several weeks | 28 volt battery voltage |
| Galileo | Satellite (navigation) | Solar cells + accumulators | 1.5 kW | > 10 years | |
| Astra 1G | Satellite (communication) | Solar cells + accumulators | 6.6 kW | > 10 years | |
| Astra 4A | Satellite (communication) | Solar cells + accumulators | 8.1 kW | > 10 years | |
| Astra 1L | Satellite (communication) | Solar cells + accumulators | 13 kW | > 10 years | |
| Meteosat-4 | Satellite (weather) | Solar cells + accumulators | 0.4 kW | 5 years | |
| SPOT-1 | Satellite (earth observation) | Solar cells + accumulators | 1.1 kW | 3 years | |
| Cassini-Huygens | Satellite (research) | RTG | 0.88 kW (0.3 kW electrical) | > 8 years | regulated 30 V DC system |
| Voyager 1 | Satellite (research) | RTG | 0.47 kW | > 12 years | regulated 30 V DC system |
| ISS | Space station | Solar cells + accumulators | 120 kW | > 12 years | 160 volts primary, 124 volts and 28 volts secondary |
Individual evidence
- ↑ a b c Ley, Wittmann, Hallmann; Space Technology Handbook; ISBN 978-3-446-41185-2
- ↑ Heinz Mielke, Transpress Lexicon: Raumfahrt - Weltraumforschung, VLN: 162-925 / 123/86
- ↑ Emcore: Space Solar Cells ( Memento from April 22, 2009 in the Internet Archive )
- ↑ Spectrolab: Solar Panel Datasheets (PDF; 190 kB)
- ↑ Bernd Leitenberger: The radioisotope elements on board space probes
- ↑ Saft-Batteries: Space-Batteries ( Memento of May 14, 2009 in the Internet Archive )
- ↑ NASA: The 1984 Goddard Space Flight Center Battery Workshop (PDF; 22.8 MB)
- ^ NASA: The NASA Aerospace Battery Safety Handbook
- ↑ Batteries and Fuel Cells in Space ( Memento from February 26, 2015 in the Internet Archive ) (PDF; 105 kB)
- ↑ SSETI ESMO Preliminary Mission / System Design Activities Spacecraft Subsystem Design Summary Description. (PDF; 92 kB) (No longer available online.) Formerly in the original ; Retrieved May 1, 2009 . ( Page no longer available , search in web archives )
- ↑ Thales Alenia Space: Electrical Power Systems ( Memento from April 19, 2016 in the Internet Archive ) (PDF, English; 5.9 MB)
- ↑ TU Delft: Space power sources (an overview) ( Memento from August 4, 2012 in the web archive archive.today )