Introduction to Astronautics Sissejuhatus kosmonautikasse

Tallinn University of Technology Introduction to AstronauticsSissejuhatus kosmonautikasse Vladislav Pustõnski 2009

Work in space is very specific since the environment is more harsh than in most of the ground applications: deep vacuum, high temperature gradients and quickly changing temperatures are the factors that technics should face to complete the mission tasks. In the previous lectures we have 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. Systems of satellites & spacecraft Power supply 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 gas generators which consume chemical propellants. Radioisotope heaters are used on some space probes and landers. Batteries & fuel cells 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 advantage of 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

Solar cells needed for short periods of time. This is the reason why they were used, for example on the Titan lander Huygens (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

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/m2 on the distance from the Sun equal to the radius of the Earth’s orbit. Efficiencies of solar cells are about 20%, so less than 300 W can 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 Earth’s 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 more expensive) gallium arsenide is also applied. To give a general idea about the performance, the value of ~100 W/kg may be adopted as a typical output per kg. Early satellites usually had fixed 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 applied today as well on satellites not exigent to high power rates. On more exigent spacecraft solar cells are placed on wings that are folded up at launch and unfold in space. Such satellites as geostationary 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 they may generate more than 10 kW. Their mass together with the frame structure and deploying

Nuclear energy sources mechanisms may be hundreds of kilograms. Attitude of spacecraft with fixed wings should be fitted with the position of the Sun, so that the angle of incidence of the solar light was larger and that the wings were not shadowed by the body of the spacecraft. In addition, the cells do not generate power than the spacecraft is in the shadow of the Earth (or, generally, of a planet), 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 independently of the attitude. However, rotation of the panels of the ISS is also used to control the drag (mostly to decrease it when the station is in the shadow, but also to increase it in order to reduce the altitude before arrival of a visiting spacecraft). Solar cells slowly degrade (about 1% per year) and their lifespan does not limit the lifespan 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 attitude corrections or reliability of other systems). However, solar flares may impact their output in higher degree. There are several types of energy sources that use nuclear reactions for energy supply in space. Most wide-spread are radioisotope thermoelectric generators (RTGs), but also radioisotope heaters and even nuclear reactors. RTGs are units that convert energy of radioisotope decay to electricity. Their design is quite simple. RTGs possess an array of thermocouples (a semiconductor device producing current when its ends have different temperatures). One side of the thermocouples is heated by

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 power distant 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 of RTGs 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

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 nuclear reactor 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 USSR Buk, 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 Jan 1978 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

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. 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 be discussed 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

the near-by celestial body (for instance, the surface of the planet for a planet orbiter); innerheat of the systems of the spacecraft. The following heat sinks are available: thermal radiation of the spacecraft; heat transfer between 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 against excessive 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 or Kevlar film sheets covered with aluminum, silver or gold, the sheets are separated 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 reflective surface 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 Giotto Halley’s comet probe was

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 as inner 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 carrier is 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 satellite PS-1. On the Lunokhods radioisotope heater was used. Sometimes the heat carrier is liquid, often water, ethylene glycol or their solutions (as on the Apollo 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 of spacesuits, but also some older spacecraft (like the Apollo LM) have been provided with such a system. Heaters may be used independently of coolant pumping systems: mostly these are electric heaters plugged into the onboard electric system and heating some assemblies, like propellant tanks, engines etc. They are switched on when needed and are cut off otherwise. Passive RHU heaters are frequently installed.

Attitude control system (ACS) Attitude control is crucial for most of spacecraft, since particular tasks of their missions (space maneuvers, orientation of instruments etc.) may be performed only with specific attitudes. Various attitude control methods are used, let us point out the basic ones. Reaction control system (RCS) Thrust is widely used on satellites and spacecraft to control their attitude. Mostly these are special attitude control engines. Actually, all methods to get thrust are used: hypergolic liquid propellant engines, electric propulsion, cold gas thrusters. For attitude control by each of 3 axes, a set of 4 thrusters is needed. One pair of the thrusters provide a positive torque and another 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 may be 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 in pulse regime, and their contribution is regulated by a number of pulses. Control thrusters together form the Reaction Control System (RCS) of the spacecraft. Besides the attitude control, the RCS may also ensure stabilization, stationkeeping, maneuvering and be a backup for operations that nominally need higher thrust, like a de-orbit maneuver and others. Engines for translations are often placed so that the vector of their thrust passes through the center of mass; thus the attitude is not altered by their firing. Otherwise they may be placed 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 & moment wheels, Control Moment Gyroscopes moment arms. Thus such engines are often installed on edges of the spacecraft or even on outriggers. Often RCS engines are redundant because of the importance of this system. They are present also on spacecraft where attitude is controlled with other methods, in this case they are used for backup and augmentation. 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 flywheels which 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 by spinning 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

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 a magnetic 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 state and 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 wheel is used not only to control the attitude, but also for stabilization purposes. Thereto it continuousty runs at very high speed (2000 – 6000 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

wheels (called RWA – Reaction Wheel Assembly) and has 2 in backup; however, work with less RWAs is possible, too. The Saturn orbiter Cassini, as well as the Messenger & 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 not rotate at all, CMGs always rotate at continuous high speeds and represent gyroscopes. In reaction & momentum wheels torque is obtained by changing the rotation speed, but in CMGs the gyroscopic moment is 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. Appearing gyroscopic moment gives 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 large spacecraft. 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 the ISS. There are two basic types of CMGs, single-gimbal and dual-gimbal. The first type provides higher torque for less energy expenses, the second type (which is used, for instance, on the ISS) 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 are needed 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 the ISS have 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

Other methods 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. Other 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 different for different parts of a spacecraft on orbit of a planet, i.e. a gravity-gradient exists: gravity field is stronger on the side oriented towards the planet. This gradient leads to appearance of tidal 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 their orbits), their influence accumulate with time. Specially high values of gravity torques are applied to elongated spacecraft. Some equilibrium positions exist for each satellite and it finally tries to settle in on of these positions, like the Moon was oriented by the gravity field of the Earth. Usually in the equilibrium position the longest axis of a satellite settles in the radial position and “looks” to the center of the planet. To orient the spacecraft with the aid of

the gravity field, it may be provided with long rods. If these rods are more than one and if their length is changeble, the equilibrium position of the satellites will be variable and it may be actively influenced by changing lengths of the rods. In the general case, when the satellite is not in the equilibrium position, it begins to oscillate around it, and to settle the satellite in the 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, the Transit constellation of navigation satellites, etc. However, gravity gradient have been used for additional stabilization of space stations, the ISS is an example. Gravity gradient is also used for desaturation of reaction wheels and CGS, it is the common practice on the ISS. Gravity-gradient stabilization is sufficiently effective only on LEO, on higher orbits the gradient is too week. However, this type of stabilization may be in perspective realized with the aid of tethers; their advantage is that tethers may be very long, so moderate weights may be placed in their ends to provide sufficient torques even on high orbits. Magnetic stabilization relays on interaction between magnets onboard the satellite and the magnetic field of a planet. A permanent magnet or a controllable electric coil inside a satellite acts as a dipole, and the magnetic field of the planet produces a torque which tries to align the axis 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-215 astronomical satellite, or to desaturate the reaction wheels or CMGs. This technique is used on the Hubble, since desaturation with the RCS might contaminate its optical surfaces. For this purpose, four 2.5-meter electronically controlled eletromagnets called Magnetic Torquers are arrayed 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.

Communication systems Aerodynamic stabilization is possible in low orbits, where drag is sufficient. Aerodynamic surfaces 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) which stabilized it attitude. Solar pressure may also be used for attitude control and desaturation purposes, specially on spacecraft with large solar panels. Each spacecraft should be able to receive commands from the Earth and to return back mission data. So spacecraft should be able to communicate with the ground (mostly a duplex data link is required), either directly or using other communication satellites for relay. Mostly radio links are used due to their universality, however, studies continue on possibility to use optical communication (via IR lasers), which is more complicated to develop and operate, but potentially offers higher data transfer rates. The tasks of spacecraft communication systems are very versatile. They include: receiving telemetry 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 commands to the spacecraft systems and the instruments onboard), relay and communication (using spacecraft 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 and the antennas. The frequency range used for communications is defined by the atmosphere

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 different spacecraft UHF is often used. Separate frequencies are used for the uplink and the downlink to allow simultaneous two-way data transmission. Since onboard power supply is generally limited, both the transmitter and the receiver should consume little power, and since communications distances are large (they may total billions of kilometers for distant planetary probes), the units should be able to work at very low signal-to-noise ratio. This is also the reason why ground antennas used to communicatewith distant probes are very large: for instance, the largest antennas of the NASA Deep Space Network (DSN) have a diameter of 70 meters (3 dishes have been built, one in California, one near Madrid and one near Canberra, Australia). However, transmitters of many communication satellites should be powerful enough so that their signal could be received by a small home antenna or even an antenna of a mobile phone. A transmitter is 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. The carrier is amplified to several tens of watts (communication satellites with a high power supply 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 to the antennas as is commanded. Once the reciever detects an uplink signal, it will follow any changes in the frequency of the signal, so that the frequency and phase of the downlink signal will be matched to the frequency and phase of the uplink. This is done with the aid of phase-lock loop (PLL). Such system is required since the uplink and downlink signals are subject to Doppler shift (which occurs due to the spacecraft proper motion, the orbital motion of the

Earth and its rotation, etc.) As a result, the signals are highly shifted from their “zero” position, and this shift is time-dependent. Without matching with the uplink signal, the downlink signal received on the ground would have a time-variable frequency that is difficult to compute. The received signal is stripped of the carrier and converted from analog into binary. Transmitters & receivers (which are always backuped) are usually combined in a single device called transponder. Often spacecraft have an array of antennas, each dedicated to a special task. So a spacecraft ordinarily has many antennas, some of them are backup and auxiliary. These antennas are mostly folded at launch and extend to the working position in space. Deep space probes are used to have large dish high gain antennas (HGAs) (with narrow beamwidth, sometimes a fraction of degree) 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 other purposes, like shielding from the Sun (the examples are the Magellan and the Cassini) and from micrometeorites (the Cassini). On the Magellan the same antenna was used for surface mapping. The HGA may be fixed or steerable, in the latest case it may be pointed independently of the attitude of the spacecraft. Communication satellites, specially GEO satellites, 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 is directed towards the incoming signal). So, the attitude of these probes and satellites should be controlled precisely, so that their antennas do not miss their targets. If there is no possibility to ensure a precise orientation, medium gain antennas (MGAs) are applied, their beamwidth is 20 – 30 degrees. The example are Lunokhods: during their drive the attitude of the body of the rover changed quickly, and a HGA would have lost the Earth, so a MGA was installed.

One member of the crew worked as “driver of the antenna”: his task was to keep the antenna pointed to the Earth. TV signals from HGAs installed on Lunar Rover Vehicles (LRVs) during the last three Apollo missions were received only on the stations, when the astronauts directed the 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 a maneuver), this operation is performed in automatic regime following a program recorded in the memory of the probe. After the operations are completed, the probe should restore the initial orientation and redirect the HGA towards the Earth. However, sometimes an error may occur: electronics may be temporarily damaged by a charged particle or some other systems may fail, and the spacecraft may loose its correct orientation. In this case communications with 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, to guarantee omnidirectional coverage). LGAs have much wider beamwidths and do not need precise pointing; however, the consequence of this property is that the strength of the signal received and transmitted by LGAs is much lower and the data rates are much smaller. If communications through the HGA fail, LGAs are used to re-establish communications, to get telemetry 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 is regained, high-rate communications are re-established. LGAs may also be used as a backup for HGAs, as it happend on the Galileo Jupiter orbiter. Its HGA failed to deploy and data transmission was performed in the low-rate regime throughout the mission.

Computers and data handling All activities of a spacecraft are controlled by the on-board computer that handles the commands from the ground and forwards them to the systems and instruments, sends back telemetry, handles data, manages fault protection and safing. However, in the early era spacecraft 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, the ISS runs on the old 80386SX CPUs. The reason is the harsh environment, 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 may be 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 taken basing on the commands of the other two). At the present time the most widespread chips are IBM RAD6000 and a more powerful version RAD750. These are radiation-hardened CPUs applied on many spacecraft and space probes. There are several subsystems in each main computer of a spacecraft. A lot of activities are performed 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. Preprogrammed actions 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 old spacecrafts these were tape recorders, at the present time they are replaced by solid state recorders, such as RAM or FLASH. Depending on the state of the spacecraft, its computer may operate the data and memory in different ways, for example, to stop scientific data collection and proceed only with the engineering data.

Safing A spacecraft should monitor itself to detect possible anomalies, faults and problems and to react 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 continuous contact 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 of fault protection algorithms are provided which are run if specific errors are detected in this or that system of the spacecraft. One of the possible responses to a serious anomaly is safing. This is a “global” procedure influencing many systems of the spacecraft. Components not urgent for functioning are shut down 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 diminish power consumption; if power supply is based on solar cells, the spacecraft tries to keep the correct 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 and payloads. Most important basic procedures for safing are permanently recorded in memory and cannot be rewritten. To increase reliability, many spacecraft systems have backup, that means, the spacecraft may continue working (in the nominal or a limited regime) if one of its systems fails.

Bus and instrumentation The major part of each spacecraft is bus. It includes the basic structure (chassis) of the spacecraft which provides places of attachment for all other systems, houses all payloads and bear all loads during the launch and maneuvers. The bus defines the geometry and mass of the spacecraft as well as its functionality. Nowadays many commercial satellites, but also other spacecraft like space probes, share a common bus. That means that there is no need to develop each spacecraft from scratch: a number of standard buses of different sizes, functionalities and prices are available on the market. Such a bus contain basic systems needed to provide functionality of a spacecraft, like tanks, ACS, RCS, computers etc. Some systems are interchangeable, for instance, smaller or larger 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 specific payloads and instrumentation (number of transmitters, their power etc.) The manufacturer adopts the bus to the client’s requirements, adding necessary systems and structures that can service the payloads and provide functionality during the needed time (solar arrays, tanks with the corresponding capacity etc.). The example is the Spacebus 4000 geostationary platform by Thales Alenia Space produced in different variants (from Spacebus 4000B2 to Spacebus 4000C4), based on the earlier Spacebus 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 kW to 16 kW) and different space avionics configurations. The masses of satellites based on this bus vary from 3 tons (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 allows to safe resources and also to increase reliability of the following probes, basing on the lessons of the previous ones. For instance, orbital modules of the Soviet Venus space probes

Life Support System Venera-9/16 were built on the same bus as the Martian probes Mars-2/6 (the joint Venus and Halley’s comet probes Vega were based on the same bus). Failed Martian probe Mars Observer was based on the Satcom-K communication satellite (structure, thermal design, solar arrays) and the TIROS/Defense meteorological satellite (attitude control, command/data handling, power conditioning, communications). Instruments and payloads placed on a spacecraft have different requirements. Some instruments, like magnetometers, are sensitive to radiation from the spacecraft. This is the reason why magnetometers are placed on extendable booms (to be as far as possible from the electric circuits of the spacecraft), RTGs are shielded in order not to influence the instruments by their heat, etc. Some instruments are placed on platforms that may be articulated independently of the body of the spacecrafts. This enables the spacecraft to maintain the desired orientation (for example, for sustaining continuous communication with the Earth) while the instruments can do their work. For instance, on the Voyager the cameras and the spectrometers are installed on such a platform. Frequently the basic systems of a spacecraft are used as scientific instruments: for instance, transmitters may be used for sounding of the atmosphere of a planet, the practice used on the Huygens Titan’s probe and others; wheels of rovers are used for soil mechanics experiments, etc. Manned spacecraft always have a Life Support System (LSS) in order to ensure normal life and 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.

On the first US manned spaceships pure O2 atmosphere was present. This was done in order 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 by fire inside the cabin during a pre-launch training. All USSR spaceships and space stations, as well as the Shuttle and the ISS, have onboard atmosphere composition similar with normal air. On the Skylab 20% of nitrogen was present, during the Apollo-Soyuz Test Project a special lock 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 the joints of the suit). During short-time flights O2 supply is provided with oxygen tanks. However, on the ISS (and on the Mir) O2 is produced from water (which is got from distilled urine) by electrolysis. If the respective device Elektron fails, O2 is produced chemically by special 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) is lowered by special dessicants. Water is delivered to orbit by cargo ships. To decrease the amount of consumables, some water is produced from distilled urine and air humidity water. Production of potable water from urine began on the ISS in 2009, as it was earlier done on the Mir. This water is also used for O2 production, hygienic and technical needs. Food is delivered by cargo ships. Temperature regime inside manned spaceships is provided by active and passive methods described above. Specially acute is this problem for spacesuits, since human body produces much heat that should be removed. Extensive active cooling is provided with water tubes passing 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 always present, and shower is provided on the ISS (as it was on the Skylab and the Mir). On the first

spaceships, also on the Apollo LM, astronauts were limited with diapers and very moderate hygienic provisions. A lot of studies have been realized concerning possibilities to create a closed (or nearly closed) LSS which would allow a continuous life support in a spaceship with minimum consumables 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 would considerably reduce the mass of a spaceship for interplanetary travels as well as would enable reduce the cost of a human base on the Moon or the Mars (since less consumables would be required to deliver from the Earth; such a delivery is very expensive and makes the base tightly dependent on deliveries). However, there have not been still created systems that could meet corresponding requirements and are sufficiently reliable.

End of the Lecture 14

Mars Reconnaissance Orbiter (By source) Solar cells on Martian spacecraft Sojourner rover (By source) Assembly of the Phoenix lander (By source)

Solar cells of ISS Solar array wing of the International Space Station (By source) ISS view during assembly (By source)

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)

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 (By source)

Reaction control system of the Apollo Position of the RCS blocks on 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)

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)

Gravity-gradient stabilization GEOS 3 gravity-gradient stabilized satellite(By source) Scheme (By Bruno Pattam, Satellite Systems)

Deep Space Network antennas 34-m Beam Waveguide antennas in Goldstone(By source) 70-m DSN antenna in Goldstone. Mass of the structure is more than 2500 tons(By source)

Voyager spacecraft Antennas Viking spacecraft Lunokhod(By source)

Mars, Venera & Vega space probes Venera-9/10 (By source) Mars-2/3 (By source) Venera-15/16 (By source) Vega-1/2(By source)

Introduction to Astronautics Sissejuhatus kosmonautikasse