Power supply for spacecraft. Power supply system for the on-board spacecraft complex (RUB 160.00) Design of the power supply system for the spacecraft
The invention relates to the field of space energy, in particular to on-board power supply systems for spacecraft (SC). According to the invention, the power supply system of a spacecraft consists of a solar battery, a voltage stabilizer, a rechargeable battery, an extreme power regulator, wherein the voltage stabilizer of the solar battery and the discharge device of the battery are made in the form of bridge inverters with a common transformer, while the input of the charger is connected to the output winding of the transformer , load power devices with their own AC or DC output voltage ratings are connected to the other output windings of the transformer, and one of the load power devices is connected to the solar battery stabilizer and the battery discharge device. The technical result is to expand the capabilities of the spacecraft power supply system, improve the quality of the output voltage, reduce development and manufacturing costs, and reduce system development time. 1 ill.
Drawings for RF patent 2396666
The present invention relates to the field of space energy, more specifically to on-board power supply systems (EPS) of spacecraft (SC).
Spacecraft power supply systems are widely known, consisting of a solar battery, a rechargeable battery, as well as a set of electronic equipment that ensures the joint operation of these sources for the spacecraft load, voltage conversion and stabilization.
Tactical and technical characteristics of the SEP, and for space technology the most important of them is specific power, i.e. the ratio of the power generated by the power supply system to its mass (Pud=Psep/Msep) depends primarily on the specific mass characteristics of the current sources used, but also to a large extent on the adopted structural diagram of the PDS, formed by the complex of electronic equipment of the PDS, which determines the modes exploitation of sources and the efficiency of using their potential.
There are known spacecraft power supply systems with structural diagrams that provide: stabilization of DC voltage on the load (with an accuracy of 0.5-1.0% of the nominal value), stabilization of voltage on the solar battery, which ensures power removal from it near the optimal operating point current-voltage characteristic (volt-ampere characteristics), and also implements optimal control algorithms for operating modes of rechargeable batteries, making it possible to ensure the highest possible capacitive parameters during long-term cycling of batteries in orbit. As an example of such power supply systems, we present the project of a power supply system for a geostationary communications spacecraft in the article A POWER, FOR A TELECOMMUNICATION SATELLITE. L.Croci, P.Galantini, C.Marana (Proceedings of the European Space Power Conference held in Graz, Austria, 23-27 August 1993 (ESA WPP-054, August 1993). Proposed PDS with a power of 5 kW, with a voltage of 42 V The efficiency of using the power of the solar battery is 97%, the efficiency of using the capacity of the battery is 80% (at the end of the 15-year service life of the spacecraft).
The structural diagram of the PDS provides for the division of the solar battery into 16 sections, each of which is regulated by its own shunt voltage stabilizer, and the outputs of the sections are connected through decoupling diodes to a common stabilized bus, which maintains 42 V ± 1%. Shunt stabilizers maintain a voltage of 42 V on the sections of the solar battery, and the design of the solar battery is carried out so that at the end of 15 years the optimal operating point of the current-voltage characteristic corresponds to this voltage.
The vast majority of foreign power supply systems and a number of domestic spacecraft, such as, for example, HS-702, A-2100 (USA), Spacebus-3000, 4000 (Western Europe), Sesat, "Express-AM", " Yamal" (Russia), etc.
In the article “Instrument complex of satellite power supply systems with extreme regulation of solar battery power”, authors V.S. Kudryashov, M.V. Nesterishin, A.V. Zhikharev, V.O. Elman, A.S. , volume 47, April 2004, No. 4) provides a description of the structural diagram of a power transmission system with an extreme solar battery power regulator, shows the effect of such regulation on the geostationary communications satellite "Express-A", which, according to the results of flight measurements, amounted to up to a 5% increase in output battery power. According to the scheme with an extreme solar battery regulator, the power supply systems of many domestic spacecraft are made, such as the geostationary spacecraft “Gals”, “Express”, high-orbit “Glonass-M”, low-orbit “Gonets”, etc.
Despite the achieved high tactical and technical characteristics of the SEP of modern spacecraft, they have a common drawback - they are not universal, which limits the scope of their use.
It is known that to power various equipment of a particular spacecraft, several ratings of the supply voltage are required, from units to tens and hundreds of volts, while in the implemented PDS a single DC power supply bus with one rating is formed, for example, 27 V, or 40 V, or 70 B, or 100 B.
When switching from one equipment supply voltage rating to another, it is necessary to develop a new power supply system with a radical redesign of current sources - solar and rechargeable batteries - and with corresponding time and financial costs.
This drawback especially affects the creation of new modifications of spacecraft based on the basic version, which is the main direction in modern spacecraft engineering.
Another disadvantage of the systems is the low noise immunity of electricity consumers on board the spacecraft. This is explained by the presence of a galvanic connection between the equipment power buses and current sources. Therefore, during sudden load fluctuations, for example, when individual consumers are switched on or off, voltage fluctuations occur on the common output bus of the power supply system, the so-called. transient processes caused by voltage surges on the internal resistance of current sources.
A power supply system with a new structural diagram is proposed, which eliminates the above-mentioned disadvantages of the known power supply systems for spacecraft.
The closest technical solution to the proposed one is the autonomous spacecraft power supply system according to RF patent 2297706, chosen as a prototype.
The prototype has the same disadvantages as the analogues discussed above.
The objective of the proposed invention is to expand the capabilities of the spacecraft power supply system, improve the quality of the output voltage, reduce development and manufacturing costs, and reduce system development time.
The essence of the claimed invention is illustrated by the drawing.
The power supply system consists of a solar battery 1, a battery 2, a solar battery voltage stabilizer 3, a battery discharge device 4, a battery charger 5, an extreme solar battery power regulator 6, connected by its inputs to discharge devices 4 and charger 5, and to a sensor. current of the solar battery 7, and the output is with a voltage stabilizer of the solar battery 3.
Stabilizer 3 and discharge device 4 are made in the form of bridge inverters. Descriptions of such bridge inverters are given, for example, in the articles: “High-frequency voltage converters with resonant switching”, author A.V. Lukin (zh. ELECTROPOPITANIE, scientific and technical collection issue 1, edited by Yu.I. Konev. Association "Power Supply" , M., 1993), The Series Connected Buck Boost Regulator For High Efficiency DC Voltage Regulation, author Arthur G. Birchenough (NASA Technical Memorandum 2003-212514, NASA Lewis Research Center, Cleveland, ON), as well as in the article BLOCK DIAGRAM AND CIRCUIT SOLUTIONS FOR AUTOMATION AND STABILIZATION COMPLEXES OF PDS OF UNSEALED GEOSTATIONARY SC WITH GALVANICA ISOLATION OF ONBOARD EQUIPMENT FROM SOLAR AND BATTERY BATTERIES authors Polyakov S.A., Chernyshev A.I., Elman V.O., Kudrya seam B.C., see "Electronic and electromechanical systems and devices: Sat. scientific works of SPC "Polyus". - Tomsk: MGP “RASKO” at the publishing house “Radio and Communications”, 2001, 568 p.
The output windings 9, 10 of the stabilizer and the discharge device are respectively connected to a common transformer 8 as its primary windings. The solar battery 1 is connected to the stabilizer 3 by plus and minus buses, and the mentioned current sensor 7 is installed in one of the buses. The battery 2 is connected to the discharge device by plus and minus buses. The charger 5 is connected by its input to the secondary winding 11 of the transformer 8, and by its output to the positive and negative buses of the battery 2.
Power devices 13 of loads 14 with their AC output voltage ratings are connected to the secondary windings 12 of transformer 8, and power devices 16 of loads 17 of DC are connected to the secondary windings 15 of transformer 8 with their voltage ratings, one of the power devices 18 of loads 19 of DC or AC , connected to the secondary winding 20 of the transformer 8, is selected as the main one, and it is used to stabilize the voltage on the secondary winding 20 of the transformer 8. For this purpose, the device 18 is connected by feedback connections to the stabilizer 3 and the discharge device 4.
The formation of an alternating voltage on the output winding 9 of the stabilizer 3 is ensured by its control circuit 21, which, according to a certain law, opens transistors 22, 23 and 24, 25 in pairs, respectively.
In a similar way, an alternating voltage is generated on the output winding of the 10-bit device 4 by its control circuit of 26 transistors 27, 28 and 29, 30, respectively.
The extreme power regulator 6, taking into account the readings of the current sensor 7 and the voltage on the solar battery 1, produces a correction signal to change the opening law of the transistors of the stabilizer 3 so that a voltage is established on the solar battery equal to the optimal voltage of the current-voltage characteristic (I-V characteristic) of the solar battery.
The power supply system operates in the following main modes.
1. Power supply of loads from a solar battery.
When the power of the solar battery exceeds the total power consumed by the loads, the bridge stabilizer 3, using the feedback of the device 18 and the stabilizer 3, on the secondary winding 20 of the transformer 8 maintains a stable voltage at a level that ensures the required voltage stability on the load 19. At the same time, on the secondary windings 11, 12, 15 of the transformer also maintain a stable alternating voltage, taking into account the transformation ratios of the windings. Battery 2 is fully charged. Charger 5 and discharge 4 are turned off, extreme regulator 6 is turned off.
2. Charge the battery.
When it becomes necessary to charge the battery, the charger 5 generates a signal to turn on the charge and provides it by converting alternating current from the secondary winding 11 of transformer 8 into direct current to charge the battery. The signal to turn on the charger 5 is also sent to the input of the extreme regulator 6, which turns on the stabilizer 3 in the extreme power control mode of the solar battery. The magnitude of the battery charging current is determined by the difference between the power of the solar battery at the optimal operating point of its current-voltage characteristics and the total power of the loads. The discharge device is disabled.
3. Power supply to the load from the battery.
This mode is formed when a spacecraft enters the shadow of the Earth or the Moon, in possible anomalous situations with loss of orientation of the solar panels, or when the spacecraft is launched into orbit when the solar panels are folded. The solar panel output is zero and the load is powered by discharging the battery. In this mode, voltage stabilization on the secondary winding 20 of transformer 8 is provided by a discharge device similar to the first mode, using feedback from device 18 to the discharge device. Stabilizer 3, extreme regulator 6, charger 5 are disabled.
4. The load is powered jointly from a solar battery and a battery.
The mode is formed when there is insufficient solar battery power to power all connected consumers, for example, when peak loads are turned on, during spacecraft maneuvers for orbit correction, during spacecraft entries and exits from shadow areas of the orbit, etc.
In this mode, the stabilizer 3 by the extreme regulator 6, following a signal from the discharge device 4, is switched on to the extreme power control mode of the solar battery 1, and the power missing to power the loads is added by discharging the battery 2. Voltage stabilization on the secondary winding 20 of the transformer 8 is provided by the discharge device 4 using feedback from device 18 to bit device 4.
The power supply system operates fully automatically.
The proposed spacecraft power supply system has the following advantages over known systems:
provides at the output the stable DC or AC voltage ratings required to power a variety of spacecraft loads, which expands its application capabilities on spacecraft of various classes or when upgrading existing devices;
higher quality of supply voltage to loads due to reduced interference, because the load power buses are galvanically (via a transformer) isolated from the current source buses;
a high degree of unification of the system is ensured and the ability to adapt it to changing conditions of use on various types of spacecraft or their modifications with minimal modification in terms of load power devices, without affecting the basic components of the system (solar and battery batteries, stabilizer, charger and discharge devices),
provides the possibility of independent design and optimization of current sources by voltage, selection of standard sizes of batteries, single solar battery generators, etc.;
The time and costs for developing and manufacturing a power supply system are reduced.
Currently at JSC "ISS" named after. M.F. Reshetnev”, together with a number of related enterprises, is developing the proposed power supply system, and manufacturing individual laboratory components of the device is underway. The first samples of the bridge inverter achieved an efficiency of 95-96.5%.
From the patent information materials known to the applicant, no set of features similar to the set of features of the claimed object was found.
CLAIM
The spacecraft power supply system, consisting of a solar battery connected by its plus and minus buses to a voltage stabilizer, a rechargeable battery connected by its plus and minus buses to the input and output of the charger, an extreme power regulator of the solar battery connected by its inputs to a current sensor, installed in one of the buses between the solar battery and the voltage stabilizer, discharge and charger devices of the battery, and the output - with the voltage stabilizer of the solar battery, characterized in that the voltage stabilizer of the solar battery and the discharge device of the battery are made in the form of bridge inverters with a common transformer, in this case, the input of the charger is connected to the output winding of the transformer, and load power devices with their own AC or DC output voltage ratings are connected to the other output windings of the transformer, and one of the load power devices is connected to the solar battery stabilizer and the battery discharge device.
Illustration copyright SPL
Space missions lasting several decades - or even longer - will require a new generation of power sources. The columnist decided to figure out what options designers have.
The power system is a vital component of a spacecraft. These systems must be extremely reliable and designed to operate under harsh conditions.
Modern complex devices require more and more energy - what does the future of their power sources look like?
The average modern smartphone can barely last a day on a single charge. And the Voyager probe, launched 38 years ago, is still transmitting signals to Earth, having already left the solar system.
Voyager's computers are capable of performing 81 thousand operations per second - but the smartphone processor works seven thousand times faster.
- Other articles on the BBC Future website in Russian
When designing a phone, of course, it is assumed that it will be regularly recharged and is unlikely to be several million kilometers from the nearest outlet.
It will not be possible to charge the battery of a spacecraft, which, according to the plan, should be located a hundred million kilometers from the current source - it needs to be able to either carry on board batteries of sufficient capacity to operate for decades, or generate electricity on its own.
It turns out that solving such a design problem is quite difficult.
Some on-board devices only need electricity occasionally, but others need to be running all the time.
Receivers and transmitters must always be turned on, and in a manned flight or on a manned space station, also life support and lighting systems.
Illustration copyright NASA Image caption The Voyager engines are not the most modern, but they have successfully served for 38 yearsDr. Rao Surampudi heads the Energy Technologies Program at the Jet Propulsion Laboratory at the California Institute of Technology in the United States. For more than 30 years, he has been developing power supply systems for various NASA vehicles.
The power system typically accounts for about 30% of a spacecraft's total mass, he said. It solves three main problems:
- power generation
- electricity storage
- electricity distribution
All of these parts of the system are vital to the operation of the device. They must weigh little, be durable and have a high “energy density” - that is, produce a lot of energy from a fairly small volume.
In addition, they must be reliable, since sending a person into space to repair breakdowns is very impractical.
The system must not only generate enough energy for all needs, but also do so throughout the entire flight - which could last for decades, and in the future, perhaps centuries.
“The design life must be long - if something breaks, there will be no one to fix it,” says Surampudi. “A flight to Jupiter takes from five to seven years, to Pluto - more than 10 years, and to leave the solar system, it takes from 20 up to 30 years old."
Illustration copyright NASA Image caption NASA's asteroid deflection mission will use a new type of solar power that is more efficient and durable than its predecessorsThe power systems of a spacecraft are subject to very specific conditions - they must remain operational in the absence of gravity, in a vacuum, under the influence of very intense radiation (which would destroy most conventional electronic devices) and extreme temperatures.
“If you land on Venus, the temperature outside will be 460 degrees,” says the specialist. “And when landing on Jupiter, the temperature will be minus 150.”
Vehicles heading towards the center of the solar system have no shortage of energy collected by their photovoltaic panels.
These panels may look a little different from solar panels installed on the roofs of residential buildings, but they operate with much higher efficiency.
It is very hot near the Sun and photovoltaic panels can overheat. To avoid this, the panels are turned away from the Sun.
In planetary orbit, photovoltaic panels are less efficient: they produce less energy, since from time to time they are fenced off from the Sun by the planet itself. In such situations, a reliable energy storage system is necessary.
Atomic solution
Such a system can be built on the basis of nickel-hydrogen batteries that can withstand more than 50 thousand charging cycles and operate for more than 15 years.
Unlike regular batteries, which don't work in space, these batteries are sealed and can function normally in a vacuum.
As you move away from the Sun, the level of solar radiation naturally decreases: for Earth it is 1374 watts per square meter, for Jupiter - 50, and for Pluto - only one watt per square meter.
Therefore, if the device flies beyond the orbit of Jupiter, then it uses atomic power systems.
The most common of these is the radioisotope thermoelectric generator (RTG), used on the Voyager, Cassini and Curiosity rover probes.
Illustration copyright NASA Image caption An improved radioisotope Stirling generator is being considered as a possible power source for long-duration missions.These power supplies have no moving parts. They produce energy from the decay of radioactive isotopes such as plutonium. Their service life exceeds 30 years.
If RTGs cannot be used (for example, if a screen that is too massive for flight is needed to protect the crew from radiation), and photovoltaic panels are not suitable because the distance from the Sun is too great, then fuel cells can be used.
Hydrogen-oxygen fuel cells were used in the American space programs Gemini and Apollo. Such cells cannot be recharged, but they release a lot of energy, and the byproduct of this process is water, which the crew can then drink.
NASA and the Jet Propulsion Laboratory are working to create more powerful, energy-intensive and compact systems with a high operating life.
But new spacecraft need more and more energy: their onboard systems are constantly becoming more complex and consume a lot of electricity.
For long flights, atomic-electric propulsion may be used
This is especially true for ships that use an electric drive - for example, ion propulsion, first used on the Deep Space 1 probe in 1998 and since then widely adopted.
Electric engines typically operate by electrically releasing fuel at high speed, but there are also those that accelerate the vehicle through electrodynamic interaction with the magnetic fields of the planets.
Most earthly energy systems are not capable of operating in space. Therefore, any new circuit undergoes a series of serious tests before being installed on a spacecraft.
NASA laboratories recreate the harsh conditions in which the new device will have to function: it is irradiated with radiation and subjected to extreme temperature changes.
Towards new frontiers
It is possible that future flights will use improved radioisotope Stirling generators. They work on a similar principle to RTGs, but are much more efficient.
In addition, they can be made very small in size - although this makes the design further complicated.
New batteries are also being created for NASA's planned flight to Europa, one of Jupiter's moons. They will be able to operate at temperatures from -80 to -100 degrees.
And the new lithium-ion batteries that designers are currently working on will have twice the capacity of the current ones. With their help, astronauts will be able, for example, to spend twice as much time on the lunar surface before returning to the ship to recharge.
Illustration copyright SPL Image caption To provide energy to such settlements, new types of fuel will most likely be required.New solar panels are also being designed that could effectively collect energy in conditions of low light and low temperatures - this will allow devices on photovoltaic panels to fly further from the Sun.
At some stage, NASA intends to establish a permanent base on Mars - and perhaps on more distant planets.
The energy systems of such settlements must be much more powerful than those currently used in space, and designed for much longer operation.
The Moon has a lot of helium-3 - this isotope is rare on Earth and is an ideal fuel for fusion power plants. However, it has not yet been possible to achieve sufficient stability of thermonuclear fusion in order to use this energy source in spacecraft.
In addition, the thermonuclear reactors that exist today occupy the space of an airplane hangar, and in this form it is impossible to use them for space flights.
Is it possible to use conventional nuclear reactors - especially in vehicles with electric propulsion and in planned missions to the Moon and Mars?
In this case, the colony will not have to maintain a separate source of electricity - the ship’s reactor can play its role.
For long flights, atomic-electric propulsion may be used.
“The Asteroid Deflection Mission requires large solar panels to provide enough electrical energy to maneuver around the asteroid,” says Surampudi. “We are currently looking at solar-electric propulsion, but nuclear-electric propulsion would be cheaper.”
However, we are unlikely to see nuclear-powered spacecraft anytime soon.
“This technology is not yet sufficiently mature. We must be absolutely sure of its safety before launching such a device into space,” explains the specialist.
Further rigorous testing is needed to ensure the reactor can withstand the rigors of spaceflight.
All of these advanced energy systems will allow spacecraft to operate longer and fly longer distances - but they are still in the early stages of development.
Once the tests are successfully completed, such systems will become a mandatory component of flights to Mars - and beyond.
- You can read it on the website.
Introduction
energy supply solar battery space
Currently, one of the priorities for the strategic development of the scientific and technical potential of the republic is the creation of the space industry. For this purpose, the National Space Agency (Kazcosmos) was created in Kazakhstan in 2007, whose activities are primarily aimed at the development and implementation of targeted space technologies and the development of space science in the interests of the socio-economic development of the country.
Scientific space research in Kazkosmos is carried out mainly at the National Center for Space Research and Technology JSC (NTSKIT JSC), which includes four research institutes: Astrophysical Institute named after. V.G. Fesenkova, Institute of Ionosphere, Institute of Space Research, Institute of Space Engineering and Technology. JSC "NTSKIT" has a large experimental base: a fleet of modern measuring equipment, test sites, observatories, scientific centers for conducting fundamental and applied scientific research in the field of space activities according to approved priorities.
Joint Stock Company "National Center for Space Research and Technology" JSC "NTSKIT" was organized through the reorganization of the Republican State Enterprise with the right of economic management "Center for Astrophysical Research" and its subsidiaries on the basis of Decree of the Government of the Republic of Kazakhstan No. 38 dated January 22, 2008.
The main subject of activity of the joint-stock company is the implementation of research, development, production and economic activities in the field of space research and technology.
One of the most important onboard systems of any spacecraft, which primarily determines its performance characteristics, reliability, service life and economic efficiency, is the power supply system. Therefore, the problems of development, research and creation of power supply systems for spacecraft are of paramount importance.
Automation of flight control processes of any spacecraft (SC) is unthinkable without electrical energy. Electrical energy is used to drive all elements of spacecraft devices and equipment (propulsion group, controls, communication systems, instrumentation, heating, etc.).
In general, the power supply system generates energy, converts and regulates it, stores it for periods of peak demand or shadow operation, and distributes it throughout the spacecraft. The power supply subsystem may also convert and regulate voltage or provide a range of voltage levels. It switches equipment on and off frequently and, to improve reliability, protects against short circuits and isolates faults. The design of the subsystem is affected by cosmic radiation, which causes degradation of solar panels. The life of a chemical battery often limits the life of a spacecraft.
Current problems are the study of the functioning features of space power sources. The study and exploration of outer space requires the development and creation of spacecraft for various purposes. Currently, automatic unmanned spacecraft are the most widely used for the formation of a global system of communications, television, navigation and geodesy, information transfer, studying weather conditions and natural resources of the Earth, as well as deep space exploration. To create them, it is necessary to ensure very stringent requirements for the accuracy of the orientation of the device in space and the correction of orbital parameters, and this requires increasing the power supply of spacecraft.
1. General information about JSC “NCIT”
Carrying out research and development work to create hardware and software for differential correction systems and consumer navigation equipment.
Object-oriented modeling and development of software and hardware for a large-scale 3D modeling system using satellite navigation technologies and laser ranging.
Development of engineering models of a complex of scientific equipment for carrying out on-board measurements and accumulating targeted scientific information and software for their operation.
Creation of scientific, methodological and software for solving problems of complex analysis and forecasting of the development of space technology in the Republic of Kazakhstan.
Creation of software and mathematical support and simulation models of spacecraft and subsystems.
Development of experimental samples of devices, equipment, components and subsystems of microsatellites.
Creation of scientific and methodological support and regulatory and technical base for solving technical regulation problems.
Regulation of requirements for the development, design, creation, operation of space technology, ensuring its safety, assessment and confirmation of compliance.
According to Government Decree No. 38 of January 22, 2008 “On the reorganization of the Republican State Enterprise “Center for Astrophysical Research” of the National Space Agency of the Republic of Kazakhstan and its subsidiary state enterprises”, the RSE “Center for Astrophysical Research” and its subsidiaries “Institute of the Ionosphere”, “Astrophysical Institute named after V.G. Fesenkov", "Institute of Space Research" were reorganized through a merger and transformation into the joint-stock company "National Center for Space Research and Technology" with 100% state participation in the authorized capital.
Certificate of state registration of JSC "NTSKIT" - No. 93168-1910-AO, identification No. 080740009161, dated July 16, 2008, registered with the Department of Justice of Almaty of the Ministry of Justice of the Republic of Kazakhstan
.2 General characteristics of the organization
Joint Stock Company "National Center for Space Research and Technology" was registered on July 16, 2008.
In the period from 2004 to July 15, 2008, JSC NTsKIT was legally the Republican State Enterprise “Center for Astrophysical Research” (with the right of economic management), which was created in accordance with the Decree of the Government of the Republic of Kazakhstan dated March 5, 2004 No. 280 “Issues some republican state enterprises of the Ministry of Education and Science of the Republic of Kazakhstan." The RSE was created on the basis of the reorganization and merger of the republican state government enterprises “Institute of Space Research”, “Institute of the Ionosphere” and “Astrophysical Institute named after V.G. Fesenkov", which were given the legal status of subsidiaries of state enterprises.
By Decree of the Government of the Republic of Kazakhstan dated May 29, 2007 No. 438 “Issues of the National Space Agency”, the RSE “Center for Astrophysical Research” (with the right of economic management) was transferred to the jurisdiction of the National Space Agency of the Republic of Kazakhstan.
The Institute of Space Research of the Academy of Sciences of the Kazakh SSR was organized in accordance with Resolution of the Cabinet of Ministers of the Kazakh SSR No. 470 of August 12, 1991. The founder and first director of the Institute is the Laureate of the State Prize of the USSR, holder of the Order of Lenin, the Red Banner of Labor, "Parasat", academician of the National Academy of Sciences of the Republic of Kazakhstan Sultangazin Umirzak Makhmutovich (1936 - 2005). In January 2011, the Institute was named after Academician U.M. Sultangazina.
The subject of the Institute's activities was conducting fundamental and applied research within the framework of state, industry, international programs and projects, as well as performing work under grants from domestic and foreign funds in the field of Earth remote sensing (ERS), space monitoring, geographic information modeling, and space materials science.
The Space Research Institute, as the parent organization, coordinated the research of the institutes of the National Academy of Sciences of the Republic of Kazakhstan and other departmental organizations in the development and implementation of all four Kazakhstani programs of scientific research and experiments on board the Mir orbital complex with the participation of cosmonaut T.O. Aubakirov. (1991) and with the participation of cosmonaut T.A. Musabaev. - (1994, 1998), on board the International Space Station - with the participation of cosmonaut T.A. Musabaev (2001).
Institute of Space Research named after academician U.M. Sultangazina was part of JSC NTsKIT as a separate legal entity in the status of a subsidiary limited liability partnership.
Since 2014The institute and the administrative apparatus of JSC "NCIT" were combined into a single structure, maintaining the personnel composition and areas of research. 1.3 Types of activities of JSC "NCIT"
Coordination, support and implementation of research activities. Fundamental and applied space research Formation of main directions and plans for scientific research, submission of completed scientific research to the National Space Agency of the Republic of Kazakhstan; Submission to the National Space Agency of the Republic of Kazakhstan of conclusions and recommendations based on annual reports of scientific organizations on scientific and scientific-technical activities; Support and implementation of experimental design and production and economic activities Creation of geographic information systems based on aerospace survey methods; Reception, processing, distribution, equivalent exchange and sale of earth remote sensing data from space; Development and operation of space assets for various purposes, space communication systems, navigation and remote sensing; Providing engineering and consulting services Conducting marketing research Implementation of innovative activities Informing about the activities of the National Space Agency - the Republic of Kazakhstan and promoting scientific achievements Propaganda of achievements of science and space technologies, organization. Conducting international and republican congresses, sessions, conferences, seminars, meetings, exhibitions; publication of scientific journals, works and information about the activities of the National Space Agency of the Republic of Kazakhstan Training of highly qualified scientific personnel. Intellectual Property Protection Development of regulatory and legal documentation Personnel composition In total - 450 qualified specialists and scientists. Among them are 27 doctors of science, 73 candidates of science, 2 academicians, 2 corresponding members and 3 PHD doctors. Center structure Department of Remote Sensing Main areas of research: Development of technologies for receiving, archiving, processing and displaying remote sensing data. Conducting fundamental and applied scientific research in the field of studying the spectral characteristics of objects on the earth's surface, space monitoring of agricultural land and the environment, emergency situations (floods, floods, fires), thematic interpretation of satellite data of various spectral, spatial and temporal resolutions based on the analysis of long-term data series Remote sensing and the state of the earth's surface. Conducting sub-satellite research. Creation of sectoral and regional situational centers for space monitoring of emergency situations. Department of Geographic Information Modeling Development of numerical models of the transfer of short-wave and thermal radiation in the atmosphere for correction of satellite images and calculations of physical parameters of the atmosphere based on satellite information. Creation of geographic information models of “risk analysis” to determine the degree of influence of natural and man-made factors on the development of emergency situations on main pipelines. Creation of automated methods and technologies for digital photogrammetry, methods and computational algorithms for interferometric analysis of remote sensing data. Department of Space Materials Science and Instrument Engineering Creation of technologies for the production of structural and functional materials for aerospace purposes, as well as products made from them. Development of qualitative, analytical and numerical methods for studying non-stationary problems in the dynamics of artificial and natural celestial bodies. Development of new mathematical models and methods for providing programmed movement of spacecraft. Department of Information and Educational Support (Astana) Organization of advanced training and retraining of specialists for the space industry of Kazakhstan. Space Information Reception Center (Almaty) and Scientific and Educational Center for Space Monitoring for Collective Use (Astana) Regular reception, archiving and processing of satellite imagery data from the Aqua/MODIS, Terra/MODIS, SuomiNPP (USA) spacecraft. There is international certification. DTOO "II" (Ionosphere Institute) Subject of activityDTOO "Institute of the Ionosphere" is conducting fundamental, exploratory and applied research in the field of solar-terrestrial physics and geodynamics: ionosphere and geomagnetic field, space weather, radiation monitoring of near-Earth space, ground-space geodynamic and geophysical monitoring of the earth's crust of Kazakhstan, creation of a forecasting system mineral deposits, geodesy and cartography. DTOO "AFIF" (Astrophysical Institute named after Fesenkov) DTOO "IKTT" (Institute of Space Engineering and Technology) Subsidiary Limited Liability Partnership "Institute of Space Engineering and Technology"(hereinafter referred to as the DTOO “Institute of Space Engineering and Technology”) was created by order of the National Space Agency of the Republic of Kazakhstan No. 65/OD dated August 17, 2009. DTOO "Institute of Space Technology and Technology" was registered on December 23, 2009. The sole founder of the Institute of Space Technology and Technology Ltd. is the National Center for Space Research and Technology Joint Stock Company. 2. General information about the power supply of spacecraft
The geometry of spacecraft, design, mass, and active life are largely determined by the power supply system of spacecraft. The power supply system or otherwise referred to as the power supply system (PSS) of spacecraft - the system of a spacecraft that provides power to other systems is one of the most important systems. Failure of the power supply system leads to failure of the entire device. The power supply system usually includes: a primary and secondary source of electricity, converters, chargers and control automation. Primary energy sources Various energy generators are used as primary sources: solar panels; chemical current sources: batteries; galvanic cells; fuel cells; radioisotope energy sources; nuclear reactors. The primary source includes not only the electricity generator itself, but also the systems that serve it, for example, the solar panel orientation system. Often energy sources are combined, for example, a solar battery with a chemical battery. Fuel cells Fuel cells have high weight and size characteristics and power density compared to a pair of solar batteries and a chemical battery, are resistant to overloads, have a stable voltage, and are silent. However, they require a supply of fuel, so they are used on devices with a period of stay in space from several days to 1-2 months. Hydrogen-oxygen fuel cells are mainly used, since hydrogen provides the highest calorific value, and, in addition, the water formed as a result of the reaction can be used on manned spacecraft. To ensure normal operation of fuel cells, it is necessary to ensure the removal of water and heat generated as a result of the reaction. Another limiting factor is the relatively high cost of liquid hydrogen and oxygen and the difficulty of storing them. Radioisotope energy sources Radioisotope energy sources are used mainly in the following cases: long flight duration; missions to the outer regions of the Solar System, where the flux of solar radiation is low; reconnaissance satellites with side-scan radar cannot use solar panels due to low orbits, but have a high energy requirement. Automation of the power supply system It includes devices for controlling the operation of the power plant, as well as monitoring its parameters. Typical tasks are: maintaining system parameters within specified ranges: voltage, temperature, pressure, switching operating modes, for example, switching to a backup power source; failure recognition, emergency protection of power supplies, in particular by current; delivery of information about the state of the system for telemetry and to the astronaut console. In some cases, it is possible to switch from automatic to manual control either from the astronaut's console or by commands from the ground control center. .1 Solar batteries operating principle and design
The solar battery is based on voltage generators made up of solar cells - devices for directly converting solar light energy into electrical energy. The action of FEP is based on the internal photoelectric effect, i.e. on the appearance of EMF under the influence of sunlight. A semiconductor photovoltaic converter (SPV) is a device that directly converts solar radiation energy into electrical energy. The operating principle of a photovoltaic cell is based on the interaction of sunlight with a semiconductor crystal, during which photons release electrons in the crystal - electrical charge carriers. Regions with a strong electric field specially created under the influence of the so-called p-n junction trap the released electrons and separate them in such a way that a current and, accordingly, electrical power arises in the load circuit. Now let's look at this process in a little more detail, albeit with significant simplifications. Let's start by looking at the absorption of light in metals and pure semiconductors. When a stream of photons hits the surface of a metal, some of the photons are reflected, and the remaining part is absorbed by the metal. The energy of the second part of the photons increases the amplitude of lattice vibrations and the speed of chaotic movement of free electrons. If the photon energy is quite high, then it may be enough to knock out an electron from the metal, giving it an energy equal to or greater than the work function of the given metal. This is an external photoelectric effect. With a lower photon energy, its energy ultimately goes entirely to heating the metal. A different picture is observed when semiconductors are exposed to a photon flux. Unlike metals, crystalline semiconductors in their pure form (without impurities), if they are not affected by any external factors (temperature, electric field, light radiation, etc.), do not have free electrons detached from the atoms of the crystal lattice of the semiconductor Rice. 2.1 - Light absorption in metals and semiconductors: 1 - filled (valence) band, 2 - band gap, 3 - conduction band, 4 - electron However, since the semiconductor material is always under the influence of some temperature (most often room temperature), a small part of the electrons can, due to thermal vibrations, acquire energy sufficient to separate them from their atoms. Such electrons become free and can take part in the transfer of electricity. A semiconductor atom that has lost an electron acquires a positive charge equal to the charge of the electron. However, a place in an atom not occupied by an electron can be occupied by an electron from a neighboring atom. In this case, the first atom becomes neutral, and the neighboring one becomes positively charged. The space vacated in an atom due to the formation of a free electron is equivalent to a positively charged particle called a hole. The energy possessed by an electron in a state bound to an atom lies within the filled (valence) band. The energy of a free electron is relatively high and lies in a higher energy band - the conduction band. Between them lies the forbidden zone, i.e. a zone of such energy values that the electrons of a given semiconductor material cannot have either in a bound or in a free state. The band gap for most semiconductors lies in the range of 0.1 - 1.5 eV. For band gap values greater than 2.0 eV, we are dealing with dielectrics. If the photon energy is equal to or exceeds the band gap, then one of the electrons is separated from its atom and transferred from the valence band to the conduction band. An increase in the concentration of electrons and holes leads to an increase in the conductivity of the semiconductor. The current conductivity in a pure single-crystal semiconductor arising under the influence of external factors is called intrinsic conductivity. With the disappearance of external influences, free electron-hole pairs recombine with each other and the intrinsic conductivity of the semiconductor tends to zero. There are no ideally pure semiconductors that have only their own conductivity. Typically, a semiconductor has electronic (n-type) or hole (p-type) conductivity. The type of conductivity is determined by the valence of the atoms of the semiconductor and the valence of the atoms of the active impurity embedded in its crystal lattice. For example, for silicon (group IV of the Mendeleev Periodic Table), the active impurities are boron, aluminum, gallium, indium, thallium (group III) or phosphorus, arsenic, antimony, bismuth (group V). The silicon crystal lattice has a shape in which each silicon atom located in a lattice site is connected to four other nearby silicon atoms by so-called covalent or pair-electronic bonds. Group V elements (donors), embedded in the sites of the silicon crystal lattice, have covalent bonds between their four electrons and the four electrons of neighboring silicon atoms, and the fifth electron can be easily released. Group III elements (acceptors), embedded in the sites of the silicon crystal lattice, attract an electron from one of the neighboring silicon atoms to form four covalent bonds, thereby forming a hole. This atom, in turn, can attract an electron from one of its neighboring silicon atoms, etc. A solar cell is a semiconductor photocell with a gate layer, the operation of which is based on the photoelectric effect just discussed. So, the mechanism of operation of the FEP is as follows (Figure 2.2). A FEP crystal consists of p- and n-regions, which have hole and electron conductivities, respectively. A p-n junction (barrier layer) is formed between these regions. Its thickness is 10-4 - 10-6 cm. Since there are more electrons on one side of the pn junction and holes on the other, each of these free current carriers will tend to diffuse into that part of the solar cell where there are not enough of them. As a result, a dynamic balance of charges is established at the p-n junction in the dark and two layers of space charges are formed, with negative charges being formed on the p-region side and positive charges on the n-region side. The established potential barrier (or contact potential difference) will prevent further self-diffusion of electrons and holes through the p-n junction. The contact potential difference Uк is directed from the n-region to the p-region. The transition of electrons from the n-region to the p-region requires the expenditure of work Uк · e, which turns into the potential energy of the electrons. For this reason, all energy levels in the p-region are raised relative to the energy levels in the n-region by the value of the potential barrier Uk · e. In the figure, the upward movement along the ordinate axis corresponds to an increase in the energy of electrons and a decrease in the energy of holes. Rice. 2.2 - Operating principle of solar cells (electrons are indicated by dots, holes are indicated by circles) Thus, the potential barrier is an obstacle for the majority carriers (in the forward direction), but does not represent any resistance for the minority carriers (in the reverse direction). Under the influence of sunlight (photons of a certain energy), the atoms of the semiconductor will be excited, and additional (excess) electron-hole pairs will appear in the crystal in both the p- and n-regions (Figure 2.2, b). The presence of a potential barrier in the p-n junction causes the separation of additional minority carriers (charges) so that excess electrons will accumulate in the n-region, and excess holes in the p-region, which did not have time to recombine before they approach the p-n junction. In this case, partial compensation of the space charge at the p-n junction will occur and the electric field created by them, directed against the contact potential difference, will increase, which together leads to a decrease in the potential barrier. As a result, a potential difference U will be established between the electrodes f , which is essentially a photo-EMF. If an external electrical load is included in the PV circuit, then an electric current will flow in it - a flow of electrons from the n-region to the p-region, where they recombine with holes. The volt-ampere and volt-power characteristics of the solar cell are presented in Figure 2.3, from which it is obvious that in order to extract maximum electrical power from the solar cell, it is necessary to ensure its operation in a fairly narrow range of output voltages (0.35 - 0.45 V). Weight 1 m 2SB 6...10 kg, of which 40% is the mass of the FEP. From photocells, the average size of which is no more than 20 mm, voltage generators are dialed by connecting them in series to the required voltage value, for example, to a nominal value of 27 V. Rice. 2.3 - Dependence of voltage and specific power on PV current density Voltage generators, having overall dimensions of approximately 100 x 150 mm, are mounted on the solar panels and connected in series to obtain the required power at the output of the solar power system. In addition to silicon solar cells, which are still used in most solar CECs, solar cells based on gallium arsenide and cadmium sulfide are of greatest interest. They have a higher operating temperature than silicon solar cells (and solar cells based on gallium arsenide have a higher theoretical and practically achieved efficiency). It should be noted that as the band gap of the semiconductor increases, the open-circuit voltage and the theoretical efficiency of a solar cell based on it increase. However, when the band gap is more than 1.5 eV, the efficiency of the solar cell begins to decrease, since an increasing proportion of photons cannot form an electron-hole pair. Thus, there is an optimal band gap (1.4 - 1.5 eV), at which the efficiency of the solar cell reaches its maximum possible value. 3. Electrochemical space power plants
An electrochemical current source (ECS) is the basis of any electrochemical CEU. It includes electrodes, which are usually active substances, an electrolyte, a separator and an external structure (vessel). An aqueous solution of KOH alkali is usually used as an electrolyte for ECHIT used on spacecraft. Let's consider a simplified diagram and design of a silver-zinc ECHIT (Figure 3.1). The positive electrode is a wire mesh current conductor onto which powdered metallic silver is pressed, then sintered in an oven at a temperature of approximately 400°C, which gives the electrode the necessary strength and porosity. The negative electrode is a mass pressed onto the current conductor grid, consisting of zinc oxide (70 - 75%) and zinc dust (25 - 30%). At the negative electrode (Zn), the oxidizing agent of the active substance reacts to zinc hydroxide Zn(OH) 2, and on the positive (AgO) - the reaction of the reduction of the active substance to pure silver. Electricity is released into the external circuit in the form of a flow of electrons. In the electrolyte, the electrical circuit is closed by the flow of OHˉ ions from the positive electrode to the negative. The separator is necessary primarily to prevent contact (and hence short circuit) of the electrodes. In addition, it reduces the self-discharge of the ECHI and is required to ensure its reversible operation over many charge-discharge cycles. Rice. 3.1 Operating principle of silver-zinc ECHIT: Positive electrode (AgO), 2 - electrical load, Negative electrode (Zn), 4 - vessel, 5 - separator The latter is due to the fact that with insufficient separation, colloidal solutions of silver oxides reaching the negative electrode are cathodically reduced in the form of thin silver threads directed towards the positive electrode, and zinc ions are also reduced in the form of threads growing towards the anode. All this can lead to a short circuit of the electrodes in the very first cycles of operation. The most suitable separator (separator) for silver-zinc ECIT is a film of hydrated cellulose (cellophane), which, swelling in the electrolyte, compacts the assembly, which prevents the zinc electrodes from melting, as well as the germination of needle-shaped silver and zinc crystals (dendrites). A silver-zinc ECHIT vessel is usually made of plastic (polyamide resin or polystyrene) and has a rectangular shape. For other types of ECHIT, vessels can be made, for example, of nickel-plated iron. When charging ECHIT, zinc and silver oxide were reduced on the electrodes. So, the ECHIT discharge is the process of releasing electricity into an external circuit, and the ECHIT charge is the process of imparting electricity to it from the outside in order to restore the original substances from the reaction products. According to the nature of their work, ECHITs are divided into galvanic cells (primary current sources), which allow only one-time use of active substances, and electric batteries (secondary current sources), which allow repeated use of active substances due to the possibility of their recovery by charging from an external source of electricity. CEUs based on ECHIT use electric batteries with disposable or reusable discharge modes, as well as hydrogen-oxygen fuel cells. 3.1 Chemical current sources
The electromotive force (EMF) of a chemical source is the difference in its electrode potentials when the external circuit is open: Where And - respectively, the potentials of the positive and negative electrodes. The total internal resistance R of a chemical source (resistance to constant current) consists of ohmic resistance and polarization resistance :
Where - EMF of polarization; - discharge current strength. Polarization resistance caused by changes in electrode potentials And when current flows and depends on the degree of charge, the strength of the discharge current, the composition of the electrodes and the purity of the electrolyte. ;
,
Where And And .
The discharge capacity Q (Ah) of a chemical source is the amount of electricity given off by the source during discharge at a certain electrolyte temperature, ambient pressure, discharge current and final discharge voltage: ,
and in the general case, with a constant current during the discharge Where - current value of discharge current, A; - discharge time, h. ,
Where And .
Silver-zinc, cadmium-nickel and nickel-hydrogen batteries are considered as chemical current sources. 3.2 Silver-zinc batteries
Silver-zinc batteries, due to their lower mass and volume with the same capacity and lower internal resistance at a given voltage, have become widespread in space electrical equipment. The active substance of the positive electrode of the battery is silver oxide AgO, and the negative plate is metallic zinc. An aqueous solution of alkali KOH with a density of 1.46 g/cm is used as an electrolyte. 3.
The battery is charged and discharged in two stages. During discharge at both stages, a zinc oxidation reaction occurs on the negative electrode 2OH ˉ discharge → ZnO + H 2O+2e. At the positive electrode, a silver reduction reaction occurs in two steps. In the first stage, divalent silver oxide is reduced to monovalent: 2AgO + 2e + H 2O discharge → Ag 2O + 2OH ˉ. The emf of the battery is 1.82.. 1.86 V. At the second stage, when the battery is discharged by approximately 30%, monovalent silver oxide is reduced to metallic silver: 2O+2e+H 2O discharge → 2Ag + 2OH ˉ. The emf of the battery at the moment of transition from the first stage of discharge to the second decreases to 1.52.. 1.56 V. As a result, curve 2 of the change in emf during discharge with the rated current (Figure 3.2) has a characteristic jump. With further discharge, the battery's emf remains constant until the battery is completely discharged. When charging, the reaction proceeds in two steps. A voltage surge and EMF occurs when the battery is approximately 30% charged (curve 1). In this state, the surface of the electrode is covered with divalent silver oxide. Rice. 3.2 - EMF of the battery during charging (1) and discharging (2) At the end of the charge, when the oxidation of silver from monovalent to divalent throughout the entire thickness of the electrode stops, the release of oxygen begins according to the equation OHˉ discharge → 2H 2O+4e+O 2
In this case, the battery emf increases by 0.2...0.3 V (see Figure 5.1, dotted section on curve 1). The oxygen released during recharging accelerates the process of destruction of the cellophane parameters of the battery and the occurrence of internal short circuits. During the charging process, all zinc oxide can be reduced to zinc metal. When recharging, the zinc oxide of the electrolyte is restored, located in the pores of the electrode, and then in the separators of the negative plates, the role of which is played by several layers of cellophane film. Zinc is released in the form of crystals that grow towards the positive electrode, forming a zinc dendrite. Such crystals can pierce cellophane films and cause short circuits of the electrodes. Zinc dendrites do not undergo reverse reactions. Therefore, even short-term overcharges are dangerous. 3.3 Nickel-cadmium batteries
The active substance of the negative electrode in a nickel-cadmium battery is cadmium metal. The electrolyte in the battery is an aqueous solution of caustic potassium KOH with a density of 1.18 ... 1.40 g/cm 3.
The nickel-cadmium battery uses a redox reaction between cadmium and nickel oxide hydrate: 2Ni(OH) 3→ Cd(OH) 2+ 2Ni(OH) 2
In a simplified way, the chemical reaction at the electrodes can be written as follows. On the negative electrode during discharge, cadmium oxidation occurs: 2e → Cd ++
Cadmium ions bind with hydroxyl ions of the alkali, forming cadmium hydrate: 2e + 2OH ˉ discharge → Cd(OH) 2.
On the positive electrode, during discharge, nickel is reduced from trivalent to divalent: 2Ni(OH) 3+ 2e discharge → 2Ni(OH)2 + 2OH ˉ. The simplification is that the composition of the hydroxide does not correspond exactly to their formulas. Cadmium and nickel salts are slightly soluble in water, so the concentration of Cd ions ++, Ni ++, Ni +++is determined by the concentration of KOH, on which the value of the battery’s emf indirectly depends in the electrolyte. The electromotive force of a newly charged battery is 1.45 V. Within a few days after the end of the charge, the EMF decreases to 1.36 V. 3.4 Nickel-hydrogen batteries
Nickel-hydrogen rechargeable batteries (HBAB), having high reliability, long service life and specific energy, and excellent performance indicators, will find wide application in spacecraft instead of nickel-cadmium batteries. To operate an LVAB in low Earth orbit (LEO), a resource of about 30 thousand cycles is required over five years. The use of batteries in LEO with a low depth of discharge (DOD) leads to a corresponding decrease in the guaranteed specific energy (30 thousand cycles can be achieved with a DOD of 40%). Three years of continuous cycling in LEO mode at GR = 30% of twelve standard NVABs (RNH-30-1) with a capacity of 30 A h showed that all NVABs operated stably for 14,600 cycles. The achieved level of specific energy for NVAB in conditions of near-Earth orbit is 40 W h/kg at a discharge depth of 100%, the resource at 30% GR is 30 thousand cycles. 4/ Selecting parameters for solar panels and buffer storage
Initial data: Limit mass of the spacecraft - MP = up to 15 kg; The height of the circular orbit is h = 450 km; The mass of the target system is no more than 0.5 kg; Transmitting frequency - 24 GHz; Voltage consumption - 3.3 - 3.6 V; The minimum power consumption of the transceiver is 300 mW; Plasma-ion engine power consumption - 155 W; The period of active existence is 2-3 years. 4.1 Calculation of buffer storage parameters
Calculation of the parameters of a buffer storage device (BN) from rechargeable batteries and determination of their composition is carried out based on the restrictions imposed on batteries in terms of charging and discharging currents, integral discharge capacity, single discharge depths, reliability, temperature operating conditions, etc. When calculating the parameters of nickel-hydrogen batteries, we will use the following characteristics and formulas [“Design of automatic spacecraft” authors: D.I. Kozlov, G.N. Anshakov, V.F. Agarkov, Yu.G. Antonov § 7.5], as well as the technical characteristics of AB HB-50 NIAI Source, information about which is taken from the site [#"justify">The electromotive force of a newly charged battery is 1.45 V. Within a few days after the end of the charge, the emf decreases up to 1.36 V. · charging current up to 30 A; · discharge current strength 12 - 50A in steady state and up to 120 A in pulsed mode for up to 1 minute; · maximum discharge depth up to 54Ah; · When operating batteries (especially in cycling modes with high charge and discharge currents), it is necessary to ensure the thermal operating conditions of the batteries in the range of 10...30°C. For this purpose, it is necessary to provide for the installation of batteries in a sealed compartment of the spacecraft and provide air cooling for each unit. Formulas used to calculate the parameters of nickel-cadmium batteries: The voltage of chemical sources of electricity differs from the EMF by the value of the voltage drop in the internal circuit, which is determined by the total internal resistance and the flowing current: , (1)
, (2)
Where And - discharge and charging voltages at the source, respectively; And - the strength of the discharge and charge currents, respectively. For galvanic cells of disposable use, the voltage is defined as discharge .
The discharge capacity Q (Ah) of a chemical source is the amount of electricity supplied by the source during discharge at a certain electrolyte temperature, ambient pressure, discharge current and final discharge voltage: , (3)
The rated capacity of a chemical current source is the capacity that the source must deliver under the operating conditions specified by the technical conditions. For KA batteries, the nominal and discharge current is most often taken to be the current of one or two or 10 hour discharge modes. Self-discharge is a useless loss of capacity by a chemical source when the external circuit is open. Typically, self-discharge is expressed as % per day of storage: (4)
Where And - chemical source containers before and after storage; T - storage time, days. The specific energy of a chemical current source is the ratio of the energy supplied to its mass: (5)
The specific energy value depends not only on the type of source, but also on the strength of the discharge current, i.e. from the taken power. Therefore, a chemical source of electricity is more fully characterized by the dependence of specific energy on specific power. Calculation of parameters: Let's determine the maximum and minimum discharge time from the formula: Therefore, the maximum discharge time is: ;
minimum discharge time: .
It follows that the discharge time allows the designed satellite to use electric current for an average of 167 minutes or 2.8 hours, since our target installation uses 89 mA, the discharge time will not be significant, which has a positive effect on the provision of electric current to other vital systems satellite Let's determine the discharge voltage and total internal resistance of the battery from the formula: ; (1)
(2)
.
From this it can be seen that the charge voltage can be sufficiently provided by using solar panels, even if they are not large in area. You can also determine self-discharge using the formula: (4)
Let's take the battery operating time T = 0.923 hours, Q 1= 50 (Ah) and Q 2 = 6 (Ah) for thirty minutes of operation: ,
that is, with a minimum current consumption of 12 A, in 30 minutes the battery will be discharged by 95% with an open circuit. Let's find the specific energy of the chemical source using the formula: ,
that is, 1 kg of chemical source can provide 61.2 W for an hour, which is also suitable for our target installation, which operates at a maximum power of 370 mW. 4.2 Calculation of solar panel parameters
To calculate the main parameters of the safety system influencing the design of the spacecraft and its technical characteristics, we will use the following formulas [“Design of automatic spacecraft” authors: D.I. Kozlov, G.N. Anshakov, V.F. Agarkov, Yu.G. Antonov § 7.5]: Calculation of SB parameters comes down to determining its area and mass. Calculation of SB power is made using the formula: (6)
Where - SB power; R n - average daily load power (without taking into account the own needs of the SEP); - time of orientation of the SB to the Sun per revolution; t T - time during which the SB is not illuminated; - The efficiency of the SB excess power regulator is 0.85; - efficiency of the BN discharge regulator equal to 0.85; R .3- efficiency of the BN charge regulator equal to 0.9; - The efficiency of BN batteries is 0.8. The area of the solar battery is calculated by the formula: (7)
Where - specific power of the SB received: W/m 2at = 60°C and 85 W/m 2at = 110°C for FEP KSP material; W/m 2at = 60°C and 100 W/m 2at = 110°C for FEP material; W/m 2at = 60°C and 160 W/m 2at = 110°C for PV material Ga - As; - safety factor, taking into account the degradation of solar cells due to radiation, equal to 1.2 for an operating time of two to three years and 1.4 for an operating time of five years; Fill factor calculated by the formula 1,12; -
SB efficiency = 0.97. The mass of the SB is determined based on specific parameters. In currently available SB designs, the specific gravity is = 2.77 kg/m 2for silicon and = 4.5 kg/m 2for gallium arsenide solar cells. The SB mass is calculated using the formula: (8)
To start calculating the PDS, you need to select solar panels. When considering various solar panels, the choice fell on the following: solar batteries of the Saturn OJSC organization based on GaAs photoconverters with the following characteristics. Basic parameters of SB Parameter of SBSB based on GaAs FPS Active lifetime, years 15 Efficiency at a temperature of 28°C, % 28 Specific power, W/m 2170Maximum power, W/m 2381Specific gravity, kg/m 21.6FEP thickness, µm150 ± 20 Also, for the calculation, you will need to know the orbital period of the satellite in low Earth orbit, information taken from the site: · in the range from 160 km the orbital period is about 88 minutes; · up to 2000 km the period is about 127 minutes. For calculation, we take the average value - about 100 minutes. At the same time, the time of illumination of the solar panels of a spacecraft in orbit is longer (about 60 minutes) than the time they are in the shadow of about 40 minutes. Load power is equal to the sum of the required power of the propulsion system, target equipment, charge power and is equal to 220 W (the value is taken with an excess of 25 W). Substituting all known values into the formula, we get: ,
.
To determine the area of the SB panel, we will take the Ga-As PV material at operating temperature = 60°C, the satellite has been operating for 2-3 years and use the formula: ,
Substituting the original data, we get: after carrying out the calculations, we get ,
but taking into account the infrequent charging of the battery, the use of modern technologies in the development of other systems, and also taking into account the fact that the load power was taken with a margin of about 25 W, it is possible to reduce the area of the power supply system to 3.6 m2
Owners of patent RU 2598862:
Usage: in the field of electrical engineering for power supply of spacecraft from primary sources of different power. The technical result is increased reliability of power supply. The power supply system of the spacecraft contains: a group of solar batteries of direct sunlight (1), a group of solar batteries of reflected sunlight (7), a generating circuit (8), a voltage stabilizer (2), a charger (3), a discharge device (4), battery (5), rectifier device (9), battery charge controller (10) and consumers (6). The alternating voltage from the generating circuit (8) is converted into constant voltage in the block (9) and is supplied to the first input of the battery charge controller (10). The constant voltage from solar panels of reflected sunlight (7) is supplied to the second input of the battery charge controller (10). The total voltage from the generating circuit and solar panels of reflected sunlight from the first output of the controller (10) goes to the second input of the battery (5). From the second output of the controller to the first input of the battery (5), control signals are received from switches (15-21) having contacts 1-3, and switches (22-25) having contacts 1-2. The number of controlled switching devices depends on the number of batteries in the battery. To recharge the selected battery (11-14) on the corresponding switches, their first contacts open with the third and close with the second, on the corresponding switches the first and second contacts close. The corresponding battery connected in this way to the second input of the battery is recharged with the rated charging current until a command is received from the controller (10) to change the next battery. The consumer (6) receives power from the remaining batteries, bypassing the disconnected one, from the first battery output (5). 5 ill.
The invention relates to space technology and can be used as part of rotation-stabilized spacecraft.
A known power supply system for a spacecraft with common buses (analogue), which contains solar panels (the primary source of energy), a battery, and consumers. The disadvantage of this system is that the voltage in this system is unstabilized. This leads to energy losses in cable networks and in built-in individual consumer stabilizers.
A known power supply system for a spacecraft with separated buses and parallel connection of a voltage stabilizer (analog), which contains a charger, a discharge device, and a battery. Its disadvantage is the impossibility of using an extreme power regulator for solar panels.
The closest in technical essence to the proposed system is a spacecraft power supply system with separated buses and with a series-parallel connection of a voltage stabilizer 2 (prototype), which also contains solar panels of direct sunlight 1, a charger 3, a discharge device 4, a rechargeable battery 5 (Fig. 1). The disadvantage of this power supply system is the inability to receive, convert and accumulate electrical energy from sources of different power, such as the energy of the Earth's magnetic field and the energy of reflected sunlight from the Earth's surface.
The purpose of the invention is to expand the capabilities of the spacecraft power supply system to receive, convert and accumulate electricity from various primary sources of different power, which allows increasing the active life and power supply of spacecraft.
In fig. 2 shows the power supply system of a rotation-stabilized spacecraft; FIG. 3 - battery containing switching devices controlled by the controller; in fig. 4 is a view of the rotation-stabilized spacecraft in FIG. Figure 5 schematically shows one of the options for the motion of a rotation-stabilized spacecraft in orbit.
The power supply system of a rotation-stabilized spacecraft contains a group of solar panels 7, designed to convert sunlight reflected from the Earth into electrical energy, generating a circuit 8, which is a set of conductors (winding) located along the body of the spacecraft, in which an electromotive force is induced for counting the rotation of the spacecraft around its axis in the Earth's magnetic field, a rectifier device 9, a battery charge controller from power sources of different power 10, a battery 5 containing controller-controlled switching devices 15-25 that connect or disconnect individual batteries 11-14 to controller 9 to recharge them with low current (Fig. 2).
The system operates as follows. During the process of launching the spacecraft into orbit, it is rotated in such a way that the axis of rotation of the apparatus and the solar panels of direct sunlight are oriented towards the Sun (Fig. 4). During the movement of a rotating spacecraft in orbit, the generating circuit intercepts the induction lines of the Earth's magnetic field at the speed of rotation of the spacecraft around its axis. As a result, according to the law of electromagnetic induction, an electromotive force is induced in the generating circuit
where µ o is the magnetic constant, H is the strength of the Earth's magnetic field, S in is the area of the generating circuit, N c is the number of turns in the circuit, ω is the angular frequency of rotation.
When the generating circuit is closed to the load, current flows in the consumer-generating circuit circuit. The power of the generating circuit depends on the torque of the spacecraft around its axis
where J KA is the moment of inertia of the spacecraft.
Thus, the generating circuit is an additional source of electricity on board the spacecraft.
The alternating voltage from the generating circuit 8 is rectified on block 9 and supplied to the first input of the battery charge controller 10. The direct voltage from the solar panels of reflected sunlight 7 is supplied to the second input of the battery charge controller 10. The total voltage from the first output of the controller 10 goes to the second input of the battery 5. From the second output of the controller to the first input of the battery 5, control signals are received from switches 15-21, having contacts 1-3, and switches 22-25, having contacts 1-2. The number of controlled switching devices depends on the number of batteries in the battery. To recharge the selected battery (11-14) on the corresponding switches, their first contacts open with the third and close with the second, on the corresponding switches the first and second contacts close. The corresponding battery connected in this way to the second input of the battery is recharged with a low current until a command is received from the controller 10 to change the next battery. The consumer receives power from the remaining batteries, bypassing battery 5, which is disconnected from the first output.
When the spacecraft is in orbit in position 1 (Fig. 4, 5), the solar panels of reflected sunlight are oriented towards the Earth. At this moment, the charger 3 included in the power supply system of the spacecraft receives electricity from solar panels of direct sunlight 1, and the battery charge controller 10 receives electricity from solar panels of reflected sunlight 7 and the generating circuit 8. In the position of the spacecraft 2, solar panels of direct solar The lights 1 remain directed towards the Sun, while the solar cells of the reflected sunlight are partially obscured. At this moment, the charger 3 of the spacecraft power supply system continues to receive electricity from solar panels of direct sunlight, and the controller 10 loses part of the energy from block 7, but continues to receive energy from block 8 through the rectifier 9. In the position of the spacecraft 3, all groups of solar panels are shaded, charger 3 does not receive electricity from solar panels 1, and on-board consumers of the spacecraft receive electricity from the battery. The battery charge controller continues to receive energy from the generating circuit 8, recharging the next battery. At the position of the spacecraft 4, the solar panels of direct sunlight 1 are again illuminated by the Sun, while the solar panels of reflected sunlight are partially obscured. At this moment, the charger 3 of the spacecraft power supply system continues to receive electricity from solar panels of direct sunlight, and the controller 10 loses some of the energy from block 7, but continues to receive energy from block 8 through the rectifier 9.
Thus, the power supply system of a rotation-stabilized spacecraft is capable of receiving, converting and accumulating: a) energy of direct and reflected from sunlight; b) kinetic energy of rotation of the spacecraft in the Earth's magnetic field. Otherwise, the functioning of the proposed system is similar to the known one.
The technical result - increasing the active life and power supply of the spacecraft - is achieved through the use of a microcontroller charger as part of the spacecraft's power supply system, which makes it possible to charge the battery from electrical energy sources of different power (reflected sunlight and energy from the Earth's magnetic field).
The practical implementation of the functional units of the present invention can be performed as follows.
A three-phase two-layer winding with an insulated copper wire can be used as a generating circuit, which will bring the shape of the electromotive force curve closer to a sinusoid. A bridge circuit of a three-phase rectifier with low-power diodes of type D2 and D9 can be used as a rectifier, which will reduce the ripple of the rectified voltage. The MAX 17710 microcontroller can be used as a battery charge controller. It can work with unstable sources with an output power range from 1 μW to 100 mW. The device has a built-in boost converter for charging batteries from sources with a typical output voltage of 0.75 V and a built-in regulator to protect batteries from overcharging. Lithium-ion batteries with a battery voltage equalization subsystem (balancing system) can be used as a battery containing controller-controlled switching devices. It can be implemented based on the MSP430F1232 controller.
Thus, the distinctive features of the proposed device contribute to achieving this goal.
Information sources
1. Analog world Maxim. New microcircuits / Symmetron Group of Companies // Issue No. 2, 2013. - 68 p.
2. Grilikhes V.A. Solar energy and space flights / V.A. Griliches, P.P. Orlov, L.B. Popov - M.: Nauka, 1984. - 211 p.
3. Kargu D.L. Power supply systems for spacecraft / D.L. Kargu, G.B. Steganov [and others] - St. Petersburg: VKA im. A.F. Mozhaisky, 2013. - 116 p.
4. Katsman M.M. Electrical machines / M.M. Katzman. - textbook manual for special students technical schools. - 2nd ed., revised. and additional - M.: Higher. Shk., 1990. - 463 p.
5. Pryanishnikov V.A. Electronics. Course of lectures / V.A. Pryanishnikov - St. Petersburg: Krona Print LLC, 1998. - 400 p.
6. Rykovanov A.N. Li-ion battery power systems / A.N. Rykovanov // Power Electronics. - 2009. - No. 1.
7. Chilin Yu.N. Modeling and optimization in spacecraft power systems / Yu.N. Chilin. - St. Petersburg: VIKA, 1995. - 277 p.
A spacecraft power supply system containing a group of solar batteries of direct sunlight, a charger that receives electricity from solar batteries of direct sunlight, a discharge device that powers consumers from a battery, a voltage stabilizer that powers consumers from a solar battery of direct sunlight, characterized in that additionally contains a group of solar panels designed to convert sunlight reflected from the Earth into electrical energy, a generating circuit, which is a set of conductors (winding) located on the body of the spacecraft, in which an electromotive force is induced due to the rotation of the spacecraft around its axis in a magnetic field the Earth field, a rectifier device, and also contains a battery charge controller from power sources of different power, a battery, which additionally contains switching devices controlled by the controller that connect or disconnect individual batteries to the controller to recharge them.
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M.A. PETROVICHEV, A. S. GURTOV SYSTEM ENERGY SUPPLY ONBOARD COMPLEX OF SPACE CARRIAGES Approved by the Editorial and Publishing Council of the University as a teaching aid SAMARA Publishing House SSAU 2007 UDC 629.78.05 BBK 39.62 P306 C T I O N A L P R E T E N A O R Y O Y E C T I O N Innovative educational program "Development of a center of competence and training of world-class specialists in the field of aerospace and geographic information technologies” PR I Reviewers: Doctor of Technical Sciences A.<...>Koptev, deputy. Head of the department of the State Scientific Research Center "TsSKB - Progress" S. I. Minenko P306 Petrovichev M.A.<...>System energy supply onboard complex spacecraft: textbook. allowance / M.A. Petrovichev, A.S. Gurtov.<...>The textbook is intended for students of specialty 160802 " Space devices and accelerating blocks."<...>UDC 629.78.05 BBK 39.62 ISBN 978-5-7883-0608-7 2 © Petrovichev M. A., Gurtov AS, 2007 © Samara State Aerospace University, 2007 System power supply on-board spacecraft complex Of all types of energy, electrical is the most universal.<...>. System power supply(SES) CA is one of the most important systems ensuring the functionality CA. <...>The reliability of SES is largely determined by 3 redundancy of all types of sources, converters, switching equipment and networks.<...>Structure systems power supply CA Basic system power supply CA is system direct current.<...>To counter load peaks use buffer source. <...>For the first time on reusable CA The Shuttle used a bufferless power supply system.<...> 4 System distribution Converter Converter Network Consumer Primary source Buffer source Rice.<...>Structure of the apparatus of the space power supply system Buffer source characterized by the fact that the total energy it produces is zero.<...>To match the characteristics of the battery with the primary source and the network, use<...>
System_of_energy_supply_of_onboard_complex_of_spacecraft_.pdf
FEDERAL AGENCY FOR EDUCATION STATE EDUCATIONAL INSTITUTION OF HIGHER PROFESSIONAL EDUCATION “SAMARA STATE AEROSPACE UNIVERSITY named after Academician S.P. QUEEN" M. A. PETROVICHEV, A. S. GURTOV POWER SUPPLY SYSTEM OF THE ON-BOARD COMPLEX OF SPACE CARRIAGES Approved by the Editorial and Publishing Council of the University as a teaching aid S A M A R A Publishing House SSAU 2007
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UDC 629.78.05 BBK 39.62 P306 Innovative educational program “Development of a center of competence and training of world-class specialists in the field of aerospace and geoinformation technologies” Reviewers: Doctor of Technical Sciences A. N. Koptev, Deputy Head of Department of the State Scientific Research Center RKTs TsSKB - Progress" S. I. M i nenko Petrovichev M. A. P306 Power supply system for the on-board complex of spacecraft: textbook / M. A. Petrovichev, A. S. Gurtov. - Samara: Samara Publishing House State Aerospace University, 2007. – 88 pp.: ill. ISBN 978-5-7883-0608-7 The role and importance of the power supply system for a spacecraft, the components of this system are considered, special attention is paid to the consideration of the principles of operation and devices of power supplies, features of their use for space technology. The manual provides quite extensive reference material that can be used in coursework and diploma design by students of non-electrical specialties. The textbook is intended for students of specialty 160802 "Spacecraft and upper stages". It may also be useful to young specialists in the rocket and space industry. Prepared at the Department of Aircraft. UDC 629.78.05 BBK 39.62 ISBN 978-5-7883-0608-7 2 © Petrovichev M. A., Gurtov AS, 2007 © Samara State Aerospace University, 2007 PRIOR I T T K E T O N E N A T I O A N L N Y P R E S
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Power supply system for the on-board spacecraft complex Of all types of energy, electrical is the most universal. Compared to other types of energy, it has a number of advantages: electrical energy is easily converted into other types of energy, the efficiency of electrical installations is much higher than the efficiency of installations operating on other types of energy, electrical energy is easy to transmit through wires to the consumer, electrical energy is easily distributed among consumers. Automation of flight control processes of any spacecraft (SC) is unthinkable without electrical energy. Electrical energy is used to drive all elements of spacecraft devices and equipment (propulsion group, controls, communication systems, instrumentation, heating, etc.). The power supply system (PSS) of a spacecraft is one of the most important systems ensuring the operation of the spacecraft. The main requirements for SES: the necessary supply of energy to complete the entire flight, reliable operation in conditions of weightlessness, the necessary reliability ensured by redundancy (in terms of power) of the main source and buffer, the absence of emissions and consumption of gases, the ability to operate in any position in space, minimal weight, minimum cost. All electrical energy necessary to carry out the flight program (for normal operation, as well as for some abnormal ones) must be on board the spacecraft, since its replenishment is possible only for manned stations. The reliability of SES is largely determined by 3
