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  • 7/30/2019 Astronautics Lecture14


    Introduction to Astronautics

    Sissejuhatus kosmonautikasse

    Vladislav Pustnski


    Tallinn University of Technology

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    Systems of satellites & spacecraftWork in space is very specific since the environment is more harsh than in most of the groundapplications: deep vacuum, high temperature gradients and quickly changing temperatures arethe factors that technics should face to complete the mission tasks. In the previous lectures wehave already described such systems as engines (including maneuvering and attitude control),tanks, guidance system and some others. Let us look other main systems of spacecraft.

    Most of spacecraft systems need energy to function. With rare exceptions the main kind of

    energy that is required by the onboard systems is electricity. The exceptions are (non-electric)

    propulsion systems and airbags for soft-landing which are powered by rocket engines or gasgenerators which consume chemical propellants. Radioisotope heaters are used on some space

    probes and landers.

    Power supply

    The first satellites (including the very first PS-1) and spacecraft (including the first space

    probes Luna-2/3, the first manned spacecraft Vostok) used batteries. The greatest advantageof batteries is their simplicity and reliability, they do not contain moving parts and do need

    special treatment (except for a suitable temperature range). At the same time a battery may

    provide very high values of voltage and current. The highest drawback of batteries is their

    relatively high weight (so the capacity-to-weight ratio is low) and a limited lifetime.

    However, they are still irreplaceable in applications where high voltages and currents are

    Batter ies & fuel cel ls

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    needed for short periods of time. This is the reason why they were used, for example on the

    Titan landerHuygens (independent lifetime the probe was totally ~3 hours of on her way to

    the Titan, during the descent in the atmosphere and on the surface). The first lunar soft-lander

    Luna-9 also was powered by batteries. Even on spacecraft with alternative power supply like

    solar cells, booster rechargeable batteries are installed for a continuous intermittent supply.

    Another possible source of electricity supply is fuel cells, these are elements that convert

    chemical energy of fuel (mostly of hydrogen and oxygen) into electricity. Conversion occurs

    in presence of a catalyst. Cells work while the fuel supply is provided. Fuel cells have an

    advantage over batteries starting from some value of required capacity (this value depends not

    only on construction but on other reasons like voltage, working temperature ranges etc.) At

    smaller required capacities, batteries have higher capacity-to-weight ratio, since they do not

    need a separate fuel tanks. Batteries are also simpler to handle and to operate with. But if high

    capacities are needed, batteries turn to be heavier than fuel cells with the fuel supply. By this

    reason fuel cells are the primary electricity supply for the Space Shuttle as they were for the

    Apollo spacecraft. The water produced by the cells as waste product was used by the crew for

    drinking and other needs. However, the lifetime of fuel cells on a spacecraft is limited by the

    fuel supply.

    Solar cells (photovoltaic cells) are the most common continuous electricity source on

    spacecraft. They convert the radiation energy of the solar light into electricity and may work

    for a very long time limited only by their aging (that in the space environment goes quicker

    Solar cells

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    because the cells are continuously degrade due to bombardment by micrometeorites and

    cosmic rays. The first spacecraft provided with solar cells was the Vanguard-1 in Mar 1958.

    The obvious advantage of solar cells that makes them irreplaceable is that they provide

    electricity supply continuously while they are illuminated by the Sun. However, they also

    have drawbacks. First, to get high power values large cells surfaces are needed. The solar

    constant is ~1370 W/m2on the distance from the Sun equal to the radius of the Earths orbit.Efficiencies of solar cells are about 20%, so less than 300 Wcan be provided by 1 m2 of a cell,

    but only if the angle of incidence is 900. At smaller angles of incidence the output is smaller.

    The output also decreases as square distance from the Sun. That means that on distant space

    probes (sent to the Mars or to the outer planets and asteroids) solar cells are much less

    effective than on the Earths orbit. So larger cells are needed for high outputs. However, solar

    cells have been used on Mars orbiters (for example, Viking, Mars Reconnaissance Orbiter

    etc.) and on Mars landers & rovers (Phoenix, Mars Exploration Rovers).

    Traditional cells are usually made from silicon, however, more efficient (but also moreexpensive) gallium arsenide is also applied. To give a general idea about the performance, thevalue of ~100 W/kgmay be adopted as a typical output per kg. Early satellites usually hadfixed solar cells, often placed on their round-shaped surface, so a part of them was illuminated

    at any attitude and another part was shadowed by the body. This simplest design is appliedtoday as well on satellites not exigent to high power rates. On more exigent spacecraft solarcells are placed on wings that are folded up at launch and unfold in space. Such satellites asgeostationary broadcast satellites need a lot of power for their work (to feed the transmitters),so the wingspan of their solar arrays may be very large, sometimes tens of meters, and theymay generate more than 10 kW. Their mass together with the frame structure and deploying

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    There are several types of energy sources that use nuclear reactions for energy supply inspace. Most wide-spread are radioisotope thermoelectric generators (RTGs), but also

    radioisotope heaters and even nuclear reactors.

    RTGs are units that convert energy ofradioisotope decay to electricity. Their design is

    quite simple. RTGs possess an array ofthermocouples (a semiconductor device producing

    current when its ends have different temperatures). One side of the thermocouples is heated by

    Nuclear energy sources

    mechanisms may be hundreds of kilograms. Attitude of spacecraft with fixed wings should befitted with the position of the Sun, so that the angle of incidence of the solar light was largerand that the wings were not shadowed by the body of the spacecraft. In addition, the cells donot generate power than the spacecraft is in the shadow of the Earth (or, generally, of aplanet), so booster batteries should be present for continuous power supply. On the

    International Space Station, as well as on other spacecraft, the solar panels track the Sun,continuously rotating so that the illumination condition of the panes are optimal independentlyof the attitude. However, rotation of the panels of theISSis also used to control the drag(mostly to decrease it when the station is in the shadow, but also to increase it in order toreduce the altitude before arrival of a visiting spacecraft).

    Solar cells slowly degrade (about 1% per year) and their lifespan does not limit thelifespan of the satellite, since they lose only ~10% - 15% of their output during the active

    lifetime of the spacecraft (which is ordinarily limited by amount of propellant for attitudecorrections or reliability of other systems). However, solar flares may impact their output inhigher degree.

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    energy released by decaying isotopes inside a sturdy container, another side is cooled by

    radiators. Most of spacecraft are powered with 238Pu RTGs. The half-life of this isotope is ~88

    years, it is mostly used as 238PuO2 oxide. Gamma and neutron radiation levels are low, which

    is very suitable for the safety reasons. Due to decay, the energy output of RTG decreases and

    thermocouples degrade; for these reasons in ~20 years an RGT loses ~20% of its initial

    capacity. However, for most spacecraft their lifespan is limited by other systems, so RGTs

    rarely limit lifespan of spacecraft. On some Soviet satellites 210Po was used with the half-life

    of ~90 days; the lifespan of these satellites was 4 months.

    The highest advantage of RTGs is their reliability: they contain no moving parts and

    nearly nothing that could fail. They are also independent of the distance from the Sun

    (contrary to photovoltaic elements) and are nearly the only energy source that can powerdistant space probes. Their disadvantage is safety: in case of a launch failure there is a

    potential hazard of disintegration of RTGs and radioactive contamination of the atmosphere

    and the zone of fall-out. This contamination is specially dangerous because of long half-life of

    the isotopes used. However, the amount of toxic elements in an RTG is usually small and their

    containers are strong enough (they are designed to survive a reentry and an impact), so the

    probability of a hazardous contamination is usually very low. In spite of this, presence ofRTGs onboard spacecraft have been criticized by some ecologists and by public.

    The first satellite with a RTG was the Transit 4A launched in Jun 1961. Since than RTGs

    have powered many Earth satellites and space probes (all launched to outer planets, like the

    Pioneer 10/11, the Voyager 1/2, etc., the Soviet Strela-1, the Martian probes Viking 1/2) as

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    well as scientific payload of the Apollo 12-17 lunar missions.

    Nuclear reactors have also been used in space: these are devices where controlled

    nuclear chain reactions are used to produce electricity. The first such reactor was onboard the

    Snapshot US experimental satellite on Apr 1965. This was the only US nuclear reactor in

    space. In USSR nuclear reactors were regularly used as powerplants of satellites. A nuclearreactor provides higher power per unit mass than a RTG, but is more complicated and less

    reliable. It is more difficult to handle and is potentially more hazardous.

    A typical nuclear powerplant consisted of a uranium reactor core and a thermoelectric

    generator (thermocouples were used in the first reactors, as the US Snap and the USSRBuk,

    but thermal emission converters were in the Topol). The mass of the Topol reactors was ~1

    ton, they contained ~11 kg of uranium fuel and had electricity output of ~5 kW (heat output

    ~150 kW). After the period of active work of the spacecraft, the reactor core was withdrawn

    by a dedicated solid rocket motor to a high-altitude orbit with the orbital decay time of ~10

    half-lives of the fuel.

    There were some incidents with the nuclear reactors in space. On May, 1968 the US

    Nimbus satellite failed to orbit, but its reactor was not destroyed and was found intact. On Jan1978 the Kosmos-954 failed to transfer the core to the graveyard orbit and fall in Canada.

    Insignificant contamination was detected (however, some very radioactive fragments were

    found), the USSR paid indemnity.

    Finally, sometimes radioisotope heater units (RHU) are used on spacecraft. Typical

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    Systems of spacecraft may normally work only in a limited range of temperatures, so the

    corresponding range should be assured for each system. Thus, the task of the thermal control

    is to protect the systems from an excessive heating and from freezing. Specific conditions of

    the outer space should be taken into account. Let us look the main issues, leaving apart the

    problems of thermal protection during descent in atmosphere (these problems will bediscussed later).

    First, vacuum is a perfect thermal insulator, so no heat conduction nor convection is

    possible between the spacecraft and the environment, only radiation transfer is possible

    (however, heat conduction is possible inside the construction of the spacecraft). The following

    heat sources should be accounted for in space:solar radiation; heat radiation by the Earth or

    Thermal control

    RHU is a small pellet of radioisotope fuel (ordinarily 238PuO2) in a strong container, natural

    decay warms them up and they keep warm the assemblies where they are attacjed. RHUs are

    used to heat up systems of spacecraft to reduce complexity (and so to increase reliability) of

    the thermal control system or if there is no other heat supply available. A typical case are deep

    space probes (specially distant from the Sun, like Martian probes and missions to the outer

    planets). The Lunokhods were provided with a 210Po RHU that heated the systems of the

    rovers during lunar nights. The RHU was placed out of the body of the rover and a reflector

    protected the body from excessive heat during the lunar days. A dedicated gas circulation

    system was installed, at nights gas passed through the RHU and blew on the systems of the

    rover. During the days the gas passage through the RHU was closed by valves.

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    the near-by celestial body (for instance, the surface of the planet for a planet orbiter); inner

    heatof the systems of the spacecraft. The following heat sinks are available: thermal

    radiation of the spacecraft; heat transferbetween the systems.

    There are two principle methods of thermal control:passive and active. Passive thermal

    control methods relay on using special thermal insulation to protect the spacecraft againstexcessive heating and also from excessive loss of heat. Another passive method is spacecraft

    spin so that it is uniformly warmed by the Sun and irradiates heat into space, and choice of a

    suitable orientation. Thermal insulation is made from multi-layer coating (multi-layer

    insulation, MTI) which generally possess high reflectance and so limits the amount of

    radiative energy absorbed and re-emitted by the spacecraft. MTI represents blankets made of

    Mylar,Kapton orKevlarfilm sheets covered with aluminum, silver or gold, the sheets areseparated by thin mesh. The working principle is as follows. Thermal equilibrium temperature

    of a surface depends on the balance between the absorbed and re-emitted heat (in the form of

    radiation), herewith the amount of heat re-emitted per second depends on the forth degree of

    the temperature, according to the Stefan-Boltzmann law. Equilibrium temperature is

    achieved when the surface re-emits all the absorbed heat. Each layer reflects back a fraction of

    the re-emitted heat, thus reducing the heat losses. At the same time, the outer reflectivesurface of the blankets reflects the incoming radiation thus protecting the spacecraft from

    overheating. Typical MTI blankets may have several tens of layers. Being made from strong

    fabrics, they also work as micrometeoroid protection, absorbing high-energy microscopic

    particles hitting the spacecraft on high velocities. (For flights near comets or planetary rings

    special micrometeoroid shields may be required, the GiottoHalleys comet probe was

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    provided with such a shield). Paints with specially chosen reflectance are also used on

    exposed parts of spacecraft. Optical Solar Reflectors (OSRs) are also applied. These are

    quartz mirror tiles which reflect solar rays and emit infrared radiation thus cooling the

    spacecraft. Such mirrors are extensively used on probes sent to inner regions of the Solar

    System, like the Magellan Venus orbiter and the Messenger Mercury orbiter. Each system

    should be accurately designed taking into account thermal conductivity of all parts as well asinner sources of heat (batteries, electric devices, RHUs etc.), equilibrium temperatures of all

    devices should be found in order to keep these temperatures inside the required ranges.

    Active thermal control methods include special systems that regulate heat exchange

    between different parts of the spacecraft. These are passages (or tubes) through which a heat

    carrieris pumped, sometimes active heaters are installed. Gaseous heat carriers are often

    used (in such case, the body of the spacecraft is pressurized, usually by pure nitrogen, and

    fans are installed to provide gas circulation inside the body). Such system was applied already

    on the very first satellitePS-1. On theLunokhods radioisotope heater was used. Sometimes

    the heat carrier is liquid, often water, ethylene glycol or their solutions (as on theApollo

    Ascent Stage). Tubes with heat carrier pass through systems providing their heating or

    cooling. Excessive heat may be ejected by radiators or by water sublimation. This is a very

    effective heat sink, but for this purpose a special water supply is needed. Most ofspacesuits,

    but also some older spacecraft (like theApollo LM) have been provided with such a system.

    Heaters may be used independently of coolant pumping systems: mostly these are electricheaters plugged into the onboard electric system and heating some assemblies, like propellanttanks, engines etc. They are switched on when needed and are cut off otherwise. Passive RHUheaters are frequently installed.

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    Attitude control system (ACS)

    Thrust is widely used on satellites and spacecraft to control their attitude. Mostly these arespecial attitude control engines. Actually, all methods to get thrust are used: hypergolicliquid propellant engines, electric propulsion, cold gas thrusters. For attitude control by eachof 3 axes, a set of 4 thrusters is needed. One pair of the thrusters provide a positive torque andanother pair, a negative torque. Thrusters in each pair are fired in the opposite directions with

    an identical thrust, providing zero net force but non-zero torque. However, these thrusters maybe fired in the same direction to provide a small net acceleration in one direction (translation).Often larger and smaller engines (verniers) are installed to allow different levels of the thrust.Chemical engines, which provide larger power, are often fired inpulse regime, and theircontribution is regulated by a number of pulses. Control thrusters together form the ReactionControl System (RCS) of the spacecraft. Besides the attitude control, the RCS may also

    ensure stabilization, stationkeeping, maneuvering and be a backup for operations thatnominally need higher thrust, like a de-orbit maneuver and others.

    Engines for translations are often placed so that the vector of their thrust passes throughthe center of mass; thus the attitude is not altered by their firing. Otherwise they may beplaced in pairs so that their combined thrust does not apply torque. Vice versa, torque-producing engines should be placed as far from the center of mass as possible for larger

    Reaction control system (RCS)

    Attitude control is crucial for most of spacecraft, since particular tasks of their missions (spacemaneuvers, orientation of instruments etc.) may be performed only with specific attitudes.Various attitude control methods are used, let us point out the basic ones.

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    Attitude control with RCS thrusters needs continuous expenditure of propellant, so the

    propellant reserve for the RCS raises the mass of a spacecraft and potentially limits its

    lifetime. However, other methods may be applied to control the attitude without spending

    propellant. The most common of them are based on moment exchange with spinning masses.

    Reaction wheels are typical attitude control devices of spacecraft. These are flywheelswhich may be spun by an electric motor and also braked down. When the motor spins the

    wheel in one direction, the stator of the motor and the body of the spacecraft (to which the

    stator is attached) start rotating in the opposite direction according to the law of conservation

    of the angular momentum. Braking down the wheel will stop the rotation of the spacecraft,

    and spinning it backwards will rotate the spacecraft back. Generally the mass (and so the

    moment of inertia) of the reaction wheel is much smaller than that of the spacecraft, so byspinning the wheel it is possible to tune very precisely the attitude. This is the reason why

    reaction wheels are often used on spacecrafts with cameras and other instruments that should

    be accurately pointed. With the use of three orthogonal wheels it is possible to organize

    attitude control by 3 axes. However, they may be combined in a single assembly, as it was

    done on the Salyut-5 space station and others. Their reaction wheels represented a heavy

    Reaction & moment wheels, Control M oment Gyroscopes

    moment arms. Thus such engines are often installed on edges of the spacecraft or even onoutriggers. Often RCS engines are redundant because of the importance of this system. Theyare present also on spacecraft where attitude is controlled with other methods, in this case theyare used for backup and augmentation.

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    sphere with a magnetic suspension, it could be spun in 3 directions with coils as a rotor in an

    electric motor.

    The quiescent state of a reaction wheel is zero rotation speed, it is spun when control

    interventions are needed. However, since starting of a wheel is accompanied by sticking in the

    rotor bearing, it is useful to run the wheel continuously at a slow speed, and/or to use amagnetic bearing of the rotor. Magnetic bearings also enable to use higher rotor speeds and

    thus to apply lighter rotors to achieve the same moment. This helps to safe weight. If a

    reaction wheel absorbs too high moment, it will rotate at the maximum speed permitted by its

    construction, and it will be unable to absorb any more moment. This phenomenon is called

    saturation. Such wheel should be desaturated. RCS thrusters may be used for this purpose,

    or other methods of moment exchange (see further): the wheel is braked to its quiescent stateand its moment is compensated by the RCS or something else.

    Moment wheel is a special type of the reaction wheel. Its specific feature is that a moment

    wheelis used not only to control the attitude, but also for stabilization purposes. Thereto it

    continuousty runs at very high speed (20006000 rmp) in the quiescent state and its quick

    rotation stabilizes the spacecraft. The stabilization principle is the same as on spinning

    spacecraft, but here only a part of the spacecraft is spun. Attitude control is realized by

    accelerating and decelerating the moment wheel, but it nominally never stops and continues

    rotating in the same direction (reaction wheels may rotate in the both directions). 3 moment

    wheels should be installed to control the attitude by 3 axes. Actually, redundant wheels are

    often used. For example, the Hubble Space Telescope normally works with 4 momentum

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    wheels (called RWAReaction Wheel Assembly) and has 2 in backup; however, work with

    less RWAs is possible, too. The Saturn orbiterCassini, as well as theMessenger& other

    spacecraft, works with 3 RWAs and has 1 in backup.

    A more complex device Control Moment Gyroscope (CMG), which uses another

    working principle. While reaction wheels usually rotate at continuous speed and may notrotate at all, CMGs always rotate at continuous high speeds and representgyroscopes. In

    reaction & momentum wheels torque is obtained by changing the rotation speed, but in CMGs

    thegyroscopic momentis used to get torque. During a control intervention, the axis of the

    wheel is inclined by actuatprs, and it results in precession of the gimbal. Appearinggyroscopic

    momentgives raise to torque, and this torque is much higher than the control intervention. So,

    a CMG works as a torque amplifier. This property make CMGs very useful on largespacecraft. This is the reason why these devices have been used on large space stations

    starting from the Skylab, later they appeared on the Mir and theISS. There are two basic

    types of CMGs,single-gimbaland dual-gimbal. The first type provides higher torque for less

    energy expenses, the second type (which is used, for instance, on theISS) enables to store

    momentum for two axes and so is more mass-efficient (since less devices are needed), but is

    more complicated and needs more power to operate with. In any case, several CMGs areneeded to provide attitude control by 3 axis as well as for redundancy reasons. Among the

    disadvantages of CMGs is their relatively high mass and large size. For instance, each of the 3

    CMGs of the Skylab had a mass of ~110 kg, their rotational frequency was ~9000 rmp. CMGs

    of theISShave a size of ~1 m and a weight of ~300 kg. However, the mass of the propellant

    which they save would be significantly larger. CMGs may also get saturated during their work

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    Other methodsOther methods applied refer rather to spacecraft stabilization when to the attitude control.They include, for example, gravity-gradient stabilization, magnetic stabilization,aerodynamic stabilization.

    Gravity-gradient stabilization is based on the fact that gravity accelerations are differentfor different parts of a spacecraft on orbit of a planet, i.e. a gravity-gradient exists: gravityfield is stronger on the side oriented towards the planet. This gradient leads to appearance oftidal forces that create torques, these torques try to rotate the spacecraft. Although the tidal

    forces are quite small (since the dimensions of spacecraft are much smaller than radii of theirorbits), their influence accumulate with time. Specially high values of gravity torques areapplied to elongated spacecraft. Some equilibrium positions exist for each satellite and itfinally tries to settle in on of these positions, like the Moon was oriented by the gravity fieldof the Earth. Usually in the equilibrium position the longest axis of a satellite settles in theradial position and looks to the center of the planet. To orient the spacecraft with the aid of

    and will need desaturation. One of the specific drawbacks of CMGs is that they have singular

    positions in their orientations (as other gimbaled devices do), so they should be operated

    carefully to avoid these singularities. Special Momentum Management Scheme (MMS) is also

    applied to use the CMGs most effectively and reduce the amount of propellant needed for the

    attitude control. Another weak point of CMGs is their reliability: they contain quickly moving

    heavy wheel, sophisticated magnetic suspension system working in vacuum and complicated

    control software. On the space stations, these are quite vulnerable devices and have been

    replaced several times.

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    the gravity field, it may be provided with long rods. If these rods are more than one and iftheir length is changeble, the equilibrium position of the satellites will be variable and it maybe actively influenced by changing lengths of the rods. In the general case, when the satelliteis not in the equilibrium position, it begins to oscillate around it, and to settle the satellite inthe equilibrium position, these oscillations should be dampened. Special dampers may be

    provided for that. A number of gravity-gradient stabilized spacecraft have been launched,among them are the GEOS geodetic satellites, the RAE 1 radio astronomy satellite, theTransit constellation of navigation satellites, etc. However, gravity gradient have been usedfor additional stabilization of space stations, theISSis an example. Gravity gradient is alsoused for desaturation of reaction wheels and CGS, it is the common practice on theISS.

    Gravity-gradient stabilization is sufficiently effective only on LEO, on higher orbits thegradient is too week. However, this type of stabilization may be in perspective realized withthe aid of tethers; their advantage is that tethers may be very long, so moderate weights maybe placed in their ends to provide sufficient torques even on high orbits.

    Magnetic stabilization relays on interaction between magnets onboard the satellite and themagnetic field of a planet. A permanent magnet or a controllable electric coil inside a satelliteacts as a dipole, and the magnetic field of the planet produces a torque which tries to align theaxis of this dipole (and thus the body of the satellite) with the lines of the field. This

    interaction may be used to de-spin a satellite, like it was done on the on the Kosmos-215astronomical satellite, or to desaturate the reaction wheels or CMGs. This technique is used ontheHubble, since desaturation with the RCS might contaminate its optical surfaces. For thispurpose, four 2.5-meter electronically controlled eletromagnets called Magnetic Torquers arearrayed along the body of the telescope. Magnetic stabilization is most effective on LEO(since the magnetic field quickly decreases with height) but was tested up to GEO.

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    Communication systemsEach spacecraft should be able to receive commands from the Earth and to return backmission data. So spacecraft should be able to communicate with the ground (mostly a duplexdata link is required), either directly or using other communication satellites for relay. Mostlyradio links are used due to their universality, however, studies continue on possibility to useoptical communication (via IR lasers), which is more complicated to develop and operate, butpotentially offers higher data transfer rates.

    The tasks of spacecraft communication systems are very versatile. They include: receivingtelemetry data (in order to track the work of the systems of a spacecraft), tracking &interchange by navigation and guidance data (in order to determine the position & attitude

    of the spacecraft and to send it to a desired route), spacecraft command (sending commandsto the spacecraft systems and the instruments onboard), relay and communication (usingspacecraft for relay), receiving mission data (collecting and sending to the ground scientific& other data).

    The basic elements of each communication system are the transmitter, the receiver andthe antennas. The frequency range used for communications is defined by the atmosphere

    Aerodynamic stabilization is possible in low orbits, where drag is sufficient. Aerodynamicsurfaces installed on the spacecraft orient it in the gas flow of the rarefied atmospheric layers.The Kosmos-149 Earth observation satellite had a round wing (truncated cone-shaped) whichstabilized it attitude.

    Solar pressure may also be used for attitude control and desaturation purposes, specially

    on spacecraft with large solar panels.

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    transparency, that is the microwave band (1 20 GHz), mostly S-band (2 4 GHz), X-band(7 12 GHz), Ku-band (12 18 GHz) (ranges are approximate). For relay between differentspacecraft UHF is often used. Separate frequencies are used for the uplink and the downlinkto allow simultaneous two-way data transmission. Since onboard power supply is generallylimited, both the transmitter and the receiver should consume little power, and sincecommunications distances are large (they may total billions of kilometers for distant planetaryprobes), the units should be able to work at very low signal-to-noise ratio. This is also thereason why ground antennas used to communicatewith distant probes are very large: forinstance, the largest antennas of the NASA Deep Space Network (DSN) have a diameter of70 meters (3 dishes have been built, one in California, one near Madrid and one nearCanberra, Australia). However, transmitters of many communication satellites should bepowerful enough so that their signal could be received by a small home antenna or even an

    antenna of a mobile phone.A transmitteris a device generating a tone at a single frequency, this tone is called the

    carrier tone. It may be sent to the Earth as a pure tone or may be modulated with data. Thecarrier is amplified to several tens of watts (communication satellites with a high powersupply may send much stronger signals) by a solid state amplifier or a traveling wave tube.The output is directed to one of the antennas (as is commanded) by waveguides.

    A receiver is sensitive to a preset narrow frequency band. It is connected by waveguides tothe antennas as is commanded. Once the reciever detects an uplink signal, it will follow anychanges in the frequency of the signal, so that the frequency and phase of the downlink signalwill be matched to the frequency and phase of the uplink. This is done with the aid ofphase-lock loop (PLL). Such system is required since the uplink and downlink signals are subject toDoppler shift (which occurs due to the spacecraft proper motion, the orbital motion of the

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    Earth and its rotation, etc.) As a result, the signals are highly shifted from their zeroposition, and this shift is time-dependent. Without matching with the uplink signal, thedownlink signal received on the ground would have a time-variable frequency that is difficultto compute. The received signal is stripped of the carrier and converted from analog intobinary. Transmitters & receivers (which are always backuped) are usually combined in asingle device called transponder.

    Often spacecraft have an array of antennas, each dedicated to a special task. So aspacecraft ordinarily has many antennas, some of them are backup and auxiliary. Theseantennas are mostly folded at launch and extend to the working position in space. Deep spaceprobes are used to have large dish high gain antennas (HGAs) (with narrow beamwidth,sometimes a fraction ofdegree) to concentrate the transmitting power in the desired direction(the Earth). This allows to use power of the transmitter effectively. The larger is the dish,

    narrower is the beamwidth of the antenna. Large HGA dishes may serve also for otherpurposes, like shielding from the Sun (the examples are theMagellan and the Cassini) andfrom micrometeorites (the Cassini). On theMagellan the same antenna was used for surfacemapping. The HGA may be fixed or steerable, in the latest case it may be pointedindependently of the attitude of the spacecraft. Communication satellites, specially GEOsatellites, frequently possess a number of HGAs directed to different regions of the Earth

    which are serviced by the specific satellite (and of course they have a receiving antenna isdirected towards the incoming signal). So, the attitude of these probes and satellites should becontrolled precisely, so that their antennas do not miss their targets. If there is no possibility toensure a precise orientation, medium gain antennas (MGAs) are applied, their beamwidth is2030 degrees. The example are Lunokhods: during their drive the attitude of the body ofthe rover changed quickly, and a HGA would have lost the Earth, so a MGA was installed.

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    One member of the crew worked as driver of the antenna: his task was to keep the antennapointed to the Earth. TV signals from HGAs installed onLunar Rover Vehicles (LRVs) duringthe last threeApollo missions were received only on the stations, when the astronauts directedthe antennas manually.

    If the attitude of a spacecraft with a fixed HGA should be intentionally changed (for

    instance, to photograph a planet during a flyby or if a probe should reorient itself to perform amaneuver), this operation is performed in automatic regime following a program recorded inthe memory of the probe. After the operations are completed, the probe should restore theinitial orientation and redirect the HGA towards the Earth. However, sometimes an error mayoccur: electronics may be temporarily damaged by a charged particle or some other systemsmay fail, and the spacecraft may loose its correct orientation. In this case communicationswith the probe should remain feasible, so that such failures would not be fatal for the mission.For this purpose low-gain antennas (LGAs) are always installed (usually more than one, toguarantee omnidirectional coverage). LGAs have much wider beamwidths and do not needprecise pointing; however, the consequence of this property is that the strength of the signalreceived and transmitted by LGAs is much lower and the data rates are much smaller. Ifcommunications through the HGA fail, LGAs are used to re-establish communications, to gettelemetry from the spacecraft, to find out the cause of the problem and to send the commands

    necessary to help the spacecraft out of the error state. When normal orientation of the HGA isregained, high-rate communications are re-established. LGAs may also be used as a backupfor HGAs, as it happend on the Galileo Jupiter orbiter. Its HGA failed to deploy and datatransmission was performed in the low-rate regime throughout the mission.

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    Computers and data handlingAll activities of a spacecraft are controlled by the on-board computer that handles thecommands from the ground and forwards them to the systems and instruments, sends backtelemetry, handles data, manages fault protection and safing. However, in the early eraspacecraft had no computers. For instance, commands to the Mercury manned spaceship

    were radioed from the ground.Actually, the computing power of the devices used in space applications is not the highest

    one. For instance, theISSruns on the old 80386SX CPUs. The reason is the harshenvironment, need for high reliability and stability against cosmic rays: a strike of a high-energy particle may crash of the system or even cause a permanent damage. The problem maybe overcome by reducing the density of transistors and by backup. In critical systems, 3

    computers may be used to run in a majority voting regime (if one fails, the decision is takenbasing on the commands of the other two). At the present time the most widespread chips areIBM RAD6000 and a more powerful version RAD750. These are radiation-hardened CPUsapplied on many spacecraft and space probes.

    There are several subsystems in each main computer of a spacecraft. A lot of activities areperformed by timer (specially on distant probes, which cannot be guided in a real time due to

    a limited speed of light), so a spacecraft clock (SCLK) always exists. Preprogrammedactions are set according to its counting and this counting is present in telemetry. Data(programs, telemetry and data to downlink) are collected on data storage devices. In oldspacecrafts these were tape recorders, at the present time they are replaced by solid staterecorders, such as RAM orFLASH. Depending on the state of the spacecraft, its computermay operate the data and memory in different ways, for example, to stop scientific datacollection and proceed only with the engineering data.

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    SafingA spacecraft should monitor itself to detect possible anomalies, faults and problems and toreact to their presence in order to mitigate a possible impact and to avoid loss of the mission.This task should be solved automatically, since the spacecraft is usually not in continuouscontact with the mission control and since the speed of light is limited, so instantaneous direct

    intervention of the mission control is mostly impossible. Thus, a number offault protectionalgorithms are provided which are run if specific errors are detected in this or that system ofthe spacecraft.

    One of the possible responses to a serious anomaly is safing. This is a global procedureinfluencing many systems of the spacecraft. Components not urgent for functioning are shutdown or reconfigured in order not to be damaged due to wrong operating. So, instrumentation

    (including scientific payloads) are switched off and programmed observations are stopped.The spacecraft may automatically try (continuously or within predefined time intervals) to re-establish pointing to the Earth and to regain communications. Measures are taken to diminishpower consumption; if power supply is based on solar cells, the spacecraft tries to keep thecorrect orientation of the solar arrays in order to avoid full exhaust of the power reserves.

    Although the nominal work of a spacecraft is disrupted when it enters into safing, strong

    and reliable protection is provided to the spacecraft as whole as well as to its systems andpayloads. Most important basic procedures for safing are permanently recorded in memoryand cannot be rewritten.

    To increase reliability, many spacecraft systems have backup, that means, the spacecraftmay continue working (in the nominal or a limited regime) if one of its systems fails.

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    Bus and instrumentationThe major part of each spacecraft is bus. It includes the basic structure (chassis) of thespacecraft which provides places of attachment for all other systems, houses all payloads andbear all loads during the launch and maneuvers. The bus defines the geometry and mass of thespacecraft as well as its functionality.

    Nowadays many commercial satellites, but also other spacecraft like space probes, share acommon bus. That means that there is no need to develop each spacecraft from scratch: anumber of standard buses of different sizes, functionalities and prices are available on themarket. Such a bus contain basic systems needed to provide functionality of a spacecraft, liketanks, ACS, RCS, computers etc. Some systems are interchangeable, for instance, smaller orlarger solar arrays may be attached depending on the requirements of the specific payload. A

    client with specific needs first chooses a bus that meets his needs and than orders specificpayloads and instrumentation (number of transmitters, their power etc.) The manufactureradopts the bus to the clients requirements, adding necessary systems and structures that canservice the payloads and provide functionality during the needed time (solar arrays, tanks withthe corresponding capacity etc.). The example is the Spacebus 4000 geostationary platformby Thales Alenia Space produced in different variants (from Spacebus 4000B2 to Spacebus4000C4), based on the earlierSpacebus 3000. It it is marketed with different heights (from

    2.2 m to 5.5 m), different output powers of its solar wings (from 8.5 kWto 16 kW) anddifferent space avionics configurations. The masses of satellites based on this bus vary from 3tons (Thor 6, based on 4000B2) to 5.7 tons (Eutelsat W2A, based on 4000C4).

    Analogically, space probes are frequently created basing on a ready bus, since that allowsto safe resources and also to increase reliability of the following probes, basing on the lessonsof the previous ones. For instance, orbital modules of the Soviet Venus space probes

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    Life Support SystemManned spacecraft always have a Life Support System (LSS) in order to ensure normal lifeand work of the crew onboard. A LSS includes supply with vital elements (oxygen, water,food) and waste product elimination (CO2, air humidity etc.), as well as hygienic resources,keeping temperature range, ventilation etc.

    Venera-9/16 were built on the same bus as the Martian probes Mars-2/6 (the joint Venus andHalleys comet probes Vega were based on the same bus). Failed Martian probe MarsObserver was based on the Satcom-Kcommunication satellite (structure, thermal design,solar arrays) and the TIROS/Defense meteorological satellite (attitude control, command/datahandling, power conditioning, communications).

    Instruments and payloads placed on a spacecraft have different requirements. Someinstruments, like magnetometers, are sensitive to radiation from the spacecraft. This is thereason why magnetometers are placed on extendable booms (to be as far as possible from theelectric circuits of the spacecraft), RTGs are shielded in order not to influence the instrumentsby their heat, etc. Some instruments are placed on platforms that may be articulatedindependently of the body of the spacecrafts. This enables the spacecraft to maintain thedesired orientation (for example, for sustaining continuous communication with the Earth)while the instruments can do their work. For instance, on the Voyagerthe cameras and thespectrometers are installed on such a platform.

    Frequently the basic systems of a spacecraft are used as scientific instruments: forinstance, transmitters may be used for sounding of the atmosphere of a planet, the practiceused on the HuygensTitans probe and others; wheels of rovers are used for soil mechanicsexperiments, etc.

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    On the first US manned spaceships pure O2 atmosphere was present. This was done inorder to decrease the inner pressure and so the mass of the structure of a spaceship. However,this required special measures to ensure fire safety. The crew of the first Apollo was killed byfire inside the cabin during a pre-launch training. All USSR spaceships and space stations, aswell as the Shuttle and theISS, have onboard atmosphere composition similar with normal air.

    On the Skylab 20% of nitrogen was present, during the Apollo-Soyuz Test Project a speciallock was provided between the spaceships. In spacesuits pure O2 atmosphere is used so far,since it enables to decrease inner pressure (with higher pressure astronaut hardly can fold thejoints of the suit). During short-time flights O2 supply is provided with oxygen tanks.However, on theISS(and on theMir) O2 is produced from water (which is got from distilledurine) by electrolysis. If the respective device Elektron fails, O2 is produced chemically byspecial cartriges. CO2 is removed from the atmosphere by a special assembly based on

    chemical absorbents. Air humidity (which tends to increase due to exhalation and sweating) islowered by special dessicants. Water is delivered to orbit by cargo ships. To decrease theamount of consumables, some water is produced from distilled urine and air humidity water.Production of potable water from urine began on theISSin 2009, as it was earlier done on theMir. This water is also used for O2 production, hygienic and technical needs. Food isdelivered by cargo ships.

    Temperature regime inside manned spaceships is provided by active and passive methodsdescribed above. Specially acute is this problem for spacesuits, since human body producesmuch heat that should be removed. Extensive active cooling is provided with water tubespassing through the inner side of suite, and the heat is ejected by sublimation.

    Hygienic needs should be also met. On the modern spaceships toilet room is alwayspresent, and shower is provided on theISS(as it was on the Skylab and theMir). On the first

    hi l h A ll li i d i h di d d

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    spaceships, also on theApollo LM, astronauts were limited with diapers and very moderatehygienic provisions.

    A lot of studies have been realized concerning possibilities to create a closed (or nearlyclosed) LSS which would allow a continuous life support in a spaceship with minimumconsumables introduced. It would include cultivation of plants (or something else) for food

    and re-cycling of waste products for water and oxygen production. Such developments wouldconsiderably reduce the mass of a spaceship for interplanetary travels as well as would enablereduce the cost of a human base on the Moon or the Mars (since less consumables would berequired to deliver from the Earth; such a delivery is very expensive and makes the basetightly dependent on deliveries). However, there have not been still created systems that couldmeet corresponding requirements and are sufficiently reliable.

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    End of the Lecture 14

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    Solar cells on Martian spacecraft

    Assembly of the Phoenix

    lander(By source)



    Orbiter(By source)


    rover(By source)
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    Solar cells of ISSSolar array wing of the

    International Space

    Station (By source)

    ISS view during

    assembly (By source)
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    Radioisotope thermoelectric generators

    RTGs of the Cassini. Shades are

    seen, they protect the probe and its

    instruments from radiative heating

    by RTGs (By source)

    Scheme of RTG of the

    Cassini (By source)
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    Multi-layer thermal insulation

    MTI on the descent stage of

    the Apollo LM (By source)New Horizons Plutonian probe

    (RTG is to the left) (By source)

    MTI blanket (Bysource)
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    Reaction control system of the Apollo

    Position of the RCS blockson the LM Ascent Stage(By source)

    One of the blocks on the CSM. 4

    identical blocks, each including 4

    engines, were installed both on the

    CSM and the LM for reduncancy(By source)
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    Reaction wheels & Control Moment Gyros

    RWA-15, used on the Swift,

    Cassini and others. 14 kg, 369

    mm diameter, 159 mm height,max. speed 2200 rmp(By source)

    Control Moment Gyroscope of

    the ISS (By source)

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    Gravity-gradient stabilization

    GEOS 3 gravity-gradientstabilized satellite(By source)

    Scheme (By Bruno Pattam, Satellite Systems)
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    Deep Space Network antennas

    70-m DSN antenna inGoldstone. Mass of the

    structure is more than 2500

    tons(By source)

    34-m Beam Waveguide

    antennas in Goldstone(By source)
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    Mars, Venera & Vega space probes

    Mars-2/3 (By source)

    V 15/16 (B ) V 1/2 (B )

    Venera-9/10 (By source)