Design of a spacecraft power supply system. Power supply system for the on-board complex of spacecraft (RUB 160.00)

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 OF AUTOMATION AND STABILIZATION COMPLEXES OF SEP OF UNSEALED GEOSTATIONARY SC WITH GALVANIAN ISOLATION OF ON-BOARD EQUIPMENT FROM SOLAR AND BATTERY BATTERIES authors Polyakov S.A., Chernyshev A.I., Elman V.O., Yashov V.S., 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, based on 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 positive and negative 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.

SOURCES OF ELECTRICAL ENERGY FOR SPACE VEHICLES
prof. Lukyanenko Mikhail Vasilievich
head Department of Automatic Control Systems of the Siberian State Aerospace University named after Academician M.F. Reshetnyova

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.
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, and their solution will allow reaching the world level in terms of specific mass indicators and active life.
Over the last decade, the world's leading companies have made a push to increase the power supply of spacecraft, which allows, with the same restrictions on the mass of the devices imposed by existing carriers, to continuously increase the power of the payload. Such achievements were made possible thanks to the efforts made by the developers of all components of on-board power supply systems, and above all, power sources.
The main sources of electricity for spacecraft currently are solar and rechargeable batteries.
Solar batteries with silicon monocrystalline photovoltaic converters have reached their physical limit in terms of mass-specific characteristics. Further progress in the development of solar cells is possible with the use of photovoltaic converters based on new materials, in particular, gallium arsenide. Three-stage photovoltaic converters made of gallium arsenide are already used on the US platform HS-702, on the European Spasebus-400, etc., which has more than doubled the power of the solar battery. Despite the higher cost of photovoltaic converters made from gallium arsenide, their use will make it possible to increase the power of a solar battery by 2-3 times or, at the same power, to correspondingly reduce the area of a solar battery compared to silicon photovoltaic converters.
Under geostationary orbit conditions, the use of photoelectric converters based on gallium arsenide makes it possible to provide a specific power of a solar battery of 302 W/m2 at the beginning of operation and 230 W/m2 at the end of its active life (10-15 years).
The development of four-stage photovoltaic converters from gallium arsenide with an efficiency of about 40% will make it possible to have a solar cell power density of up to 460 W/m2 at the beginning of operation and 370 W/m2 at the end of its active life. In the near future, we should expect a significant improvement in the mass-specific characteristics of solar cells.
Currently, batteries based on the nickel-hydrogen electrochemical system are widely used on spacecraft; however, the energy-mass characteristics of these batteries have reached their limit (70-80 Wh/kg). The possibility of further improving the specific characteristics of nickel-hydrogen batteries is very limited and requires large financial costs.
To create competitive space technology, it was necessary to switch to new types of electrochemical power sources suitable for use as part of the power supply system for promising spacecraft.
The space technology market is currently actively introducing lithium-ion batteries. This is because lithium-ion batteries have a higher energy density compared to nickel-hydrogen batteries.
The main advantage of the lithium-ion battery is the reduction in weight due to the higher energy-to-weight ratio. The energy-weight ratio of lithium-ion batteries is higher (125 Wh/kg) compared to the maximum achieved for nickel-hydrogen batteries (80 Wh/kg).
The main advantages of lithium-ion batteries are:
- reduction in battery weight due to a higher energy-to-weight ratio (weight reduction for the battery is ~40%);
- low heat generation and high energy efficiency (charge-discharge cycle) with very low self-discharge, which ensures the simplest control during launch, transfer orbit and normal operation;
- a more technologically advanced manufacturing process for lithium-ion batteries compared to nickel-hydrogen batteries, which allows for good repeatability of characteristics, high reliability and reduced costs.
According to experts from SAFT (France), the use of lithium-ion batteries on telecommunications satellites with a power of 15-20 kW will reduce the mass of batteries by 300 kg (the cost of putting 1 kg of useful mass into orbit is ~$30,000).
Main characteristics of the VES140 lithium-ion battery (developed by SAFT): guaranteed capacity 39 A*h, average voltage 3.6 V, end-of-charge voltage 4.1 V, energy 140 Wh, specific energy 126 Wh/kg , weight 1.11 kg, height 250 mm and diameter 54 mm. The VES140 battery is qualified for space applications.
In Russia, today OJSC Saturn (Krasnodar) has developed and manufactured the lithium-ion battery LIGP-120. Main characteristics of the LIGP-120 battery: nominal capacity 120 Ah, average voltage 3.64 V, specific energy 160 Wh/kg, weight 2.95 kg, height 260 mm, width 104.6 mm and depth 44.1 mm. The battery has a prismatic shape, which provides significant advantages in terms of specific volumetric energy of the battery compared to SAFT batteries. By varying the geometric dimensions of the electrode, you can obtain a battery of different capacities. This design provides the highest specific-volume characteristics of the battery and allows the battery to be configured, ensuring optimal thermal conditions.
Modern power supply systems for spacecraft are a complex complex of power sources, converting and distribution devices, integrated into an automatic control system and designed to power on-board loads. Secondary power supplies are an energy-converting complex consisting of a certain number of identical pulse voltage converters operating for a common load. In the traditional version, classical converters with a rectangular shape of the current and voltage of the key element and control via pulse width modulation are used as pulse voltage converters.
To improve the technical and economic indicators of the spacecraft power supply system, such as power density, efficiency, speed, and electromagnetic compatibility, we proposed the use of quasi-resonant voltage converters. Studies were carried out on the operating modes of two parallel-connected quasi-resonant serial-type voltage converters with switching of an electronic switch at zero current values and a pulse-frequency control law. Based on the results of modeling and studying the characteristics of prototypes of quasi-resonant voltage converters, the advantages of this type of converters were confirmed.
The results obtained allow us to conclude that the proposed quasi-resonant voltage converters will find wide application in power supply systems for digital and telecommunication systems, instrumentation, process equipment, automation and telemechanics systems, security systems, etc.
Current problems are the study of the functioning features of space power sources, the development of their mathematical models and the study of energy and dynamic regimes.
For these purposes, we have developed and manufactured unique equipment for studying power supply systems of spacecraft, which allows automated testing of on-board power sources (solar and rechargeable batteries) and power supply systems in general.
In addition, an automated workstation for studying the energy-thermal conditions of lithium-ion batteries and battery modules and a hardware complex for studying the energy and dynamic characteristics of gallium arsenide solar cells were developed and manufactured.
An important aspect of the work is also the creation and research of alternative sources of electricity for spacecraft. We have conducted research on a flywheel energy storage device, which is a super flywheel combined with an electric machine. A flywheel rotating in a vacuum on magnetic supports has an efficiency of 100%. The two-rotor flywheel energy storage device has a property that makes it possible to realize a triaxial angular orientation. In this case, the power gyroscope (gyrodine), as an independent separate subsystem, can be excluded, i.e. The flywheel energy storage device combines the functions of an energy storage device and a power gyroscope.
Research has been carried out on electrodynamic tether systems as a source of electricity for a spacecraft. To date, a mathematical model of an electrodynamic cable system has been developed to calculate maximum power; the dependences of energy characteristics on orbital parameters and tether length were determined; a methodology has been developed for determining the parameters of a cable system that ensures the generation of a given power; the orbital parameters (height and inclination) at which the most efficient use of tether systems in energy generation mode is achieved are determined; The capabilities of the cable system when operating in traction mode were investigated.

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 inenko Petrovichev P306 Power supply system for on-board spacecraft complex: textbook / M.A. Petrovichev, A.S. Gurtov. State Aerospace University, 2007. – 88 p.: ill. 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. 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

6 solar panels are clearly visible, rigidly fixed to the body. To maximize the power of such an installation, a constant orientation of the apparatus body to the Sun is necessary, which required the development of an original orientation control system

Spacecraft power supply system (power supply system, BOT) - the spacecraft system, which provides power to other systems, is one of the most important systems; in many ways, it determines the geometry of spacecraft, design, mass, and active life. 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.

System parameters

The required power of the apparatus's power plant is constantly growing as new tasks are mastered. Thus, the first artificial Earth satellite (1957) had a power plant with a power of about 40 W, the Molniya-1+ device (1967) had a power plant of 460 W, the communications satellite Yakhsat 1B (2011) - 12 kW.

Today, most of the onboard equipment of foreign-made spacecraft is powered by a constant voltage of 50 or 100 volts. If it is necessary to provide the consumer with alternating voltage or constant voltage of a non-standard value, static semiconductor converters are used.

Primary energy sources

Various energy generators are used as primary sources:

, in particular:

The primary source includes not only the electricity generator itself, but also the systems that serve it, for example, a solar panel orientation system.

Often energy sources are combined, for example, a solar battery with a chemical battery.

Solar panels

Today, solar panels are considered one of the most reliable and well-proven options for providing energy to a spacecraft.

The radiation power of the Sun in Earth's orbit is 1367 W/m². This allows you to receive approximately 130 W per 1 m² of solar panel surface (with an efficiency of 8...13%). Solar panels are located either on the outer surface of the device or on drop-down rigid panels. To maximize the energy released by the batteries, the perpendicular to their surface should be directed towards the Sun with an accuracy of 10...15˚. In the case of rigid panels, this is achieved either by the orientation of the spacecraft itself or by a specialized autonomous electromechanical system for orienting solar panels, while the panels are movable relative to the body of the device. Some satellites use non-orientable batteries, placing them on the surface so that the required power is provided at any position of the device.

Solar panels degrade over time due to the following factors:

meteor erosion reducing the optical properties of the surface of photoelectric converters;
radiation that lowers photovoltage, especially during solar flares and when flying in the Earth's radiation belt;
thermal shocks due to deep cooling of the structure in the shaded areas of the orbit, heating in the illuminated areas and vice versa. This phenomenon destroys the fastening of individual battery elements and the connections between them.

There are a number of measures to protect batteries from these phenomena. The effective operation time of solar batteries is several years; this is one of the limiting factors that determine the active life of a spacecraft.

When the batteries are shaded as a result of maneuvers or entering the shadow of a planet, energy production by photoelectric converters stops, so the power supply system is supplemented with chemical batteries (chemical buffer batteries).

Rechargeable batteries

The most common in space technology are nickel-cadmium batteries, as they provide the largest number of charge-discharge cycles and have better overcharge resistance. These factors come to the fore when the device has a service life of more than a year. Another important characteristic of a chemical battery is specific energy, which determines the weight and size characteristics of the battery. Another important characteristic is reliability, since redundancy of chemical batteries is highly undesirable due to their high mass. Batteries used in space technology are usually sealed; tightness is usually achieved using metal-ceramic seals. The batteries also have the following requirements:

high specific weight and size characteristics;
high electrical characteristics;
wide range of operating temperatures;
possibility of charging with low currents;
low self-discharge currents.

In addition to its main function, the battery can play the role of a voltage stabilizer for the on-board network, since in the operating temperature range its voltage changes little when the load current changes.

Fuel cells

This type of power source was first used on the Gemini spacecraft in 1966. Fuel cells have high weight-dimensional 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.

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 powers (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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