System for controlling the position of solar panels of a spacecraft. Method for controlling the position of solar panels of a spacecraft and a system for its implementation Sensors for moving solar panels of spacecraft

The Roman philosopher Seneca said: “If a person does not know where he is sailing, then there is no favorable wind for him.” In fact, what use is it to us if we do not know the position of the device in space? This story is about devices that allow us not to get lost in space.

Technological advances have made attitude control systems small, cheap and accessible. Now even a student microsatellite can boast an orientation system that the pioneers of astronautics could only dream of. Limited opportunities gave rise to ingenious solutions.

Asymmetrical answer: no orientation

The first satellites and even interplanetary stations flew unoriented. Data transmission to Earth was carried out via a radio channel, and several antennas, so that the satellite could be in touch at any position and any tumbles, weighed much less than the attitude control system. Even the first interplanetary stations flew unoriented:

Luna 2, the first station to reach the lunar surface. Four antennas on the sides provide communication at any position relative to the Earth

Even today, it is sometimes easier to cover the entire surface of a satellite with solar panels and install several antennas than to create an attitude control system. Moreover, some tasks do not require orientation - for example, cosmic rays can be detected in any position of the satellite.

Advantages:

• Maximum simplicity and reliability. A missing orientation system cannot fail.

Flaws:

• Currently suitable mainly for microsatellites that solve relatively simple problems. “Serious” satellites can no longer do without an attitude control system.

Solar sensor

By the middle of the 20th century, photocells had become a familiar and mastered thing, so it is not surprising that they went into space. The Sun became an obvious beacon for such sensors. Its bright light fell on the photosensitive element and made it possible to determine the direction:

Various operation schemes of modern solar sensors, at the bottom there is a photosensitive matrix

Another design option, here the matrix is ​​curved

Modern solar sensors

Advantages:

• Simplicity.


• Cheapness.


• The higher the orbit, the smaller the shadow area, and the longer the sensor can operate.


• The accuracy is approximately one arc minute.

Flaws:

• Do not work in the shadow of the Earth or other celestial body.


• May be subject to interference from the Earth, Moon, etc.

Just one axis along which solar sensors can stabilize the device does not interfere with their active use. Firstly, the solar sensor can be complemented with other sensors. Secondly, for spacecraft with solar batteries, the solar sensor makes it easy to organize a rotation mode on the Sun, when the device rotates aimed at it, and the solar batteries operate in the most comfortable conditions.
The Vostok spacecraft cleverly used a solar sensor - the axis on the Sun was used when constructing the orientation to decelerate the ship. Also, solar sensors were in great demand on interplanetary stations, because many other types of sensors cannot operate outside of Earth orbit.
Due to their simplicity and low cost, solar sensors are now very common in space technology.

Infrared vertical

Vehicles that fly in Earth orbit often need to determine the local vertical - the direction towards the center of the Earth. Visible photocells are not very suitable for this - on the night side the Earth is much less illuminated. But, fortunately, in the infrared range, the warm Earth shines almost equally on the day and night hemispheres. In low orbits, sensors determine the position of the horizon; in high orbits, they scan the space in search of the warm circle of the Earth.
Structurally, as a rule, infrared vertical plotters contain a system of mirrors or a scanning mirror:

Infrared vertical assembly with flywheel. The unit is designed for precise orientation to the Earth for geostationary satellites. The scanning mirror is clearly visible

An example of the field of view of the infrared vertical. Black circle - Earth

Domestic infrared verticals produced by JSC "VNIIEM"

Advantages:

• Capable of building a local vertical in any part of the orbit.


• Generally high reliability.


• Good accuracy -

Flaws:

• Orientation on one axis only.


• For low orbits, certain designs are needed, for high orbits, others.


• Relatively large dimensions and weight.


• Only for Earth orbit.

The fact that the orientation is constructed along only one axis does not prevent the widespread use of infrared verticals. They are very useful for geostationary satellites that need to point their antennas towards the Earth. ICRs are also used in manned cosmonautics, for example, on modern modifications of the Soyuz spacecraft, orientation to braking is carried out only according to its data:

The Soyuz ship. Duplicate SCI sensors are shown by arrows

Gyroorbitant

In order to issue a braking impulse, it is necessary to know the direction of the orbital velocity vector. The solar sensor will give the correct axis approximately once a day. This is normal for astronaut flights; in case of an emergency, a person can manually orient the ship. But the Vostok ships had “twin brothers”, the Zenit reconnaissance satellites, which also needed to issue a braking impulse in order to return the captured film from orbit. The limitations of the solar sensor were unacceptable, so something new had to be invented. This solution was the gyroorbitant. When the infrared vertical works, the ship rotates because the axis to Earth is constantly turning. The direction of orbital motion is known, so by the direction in which the ship turns, its position can be determined:

For example, if the ship constantly rolls to the right, then we are flying right side forward. And if the ship flies stern forward, then it will constantly raise its nose up. With the help of a gyroscope, which tends to maintain its position, this rotation can be determined:

The more the arrow is deflected, the more pronounced the rotation along this axis. Three such frames allow you to measure rotation along three axes and turn the ship accordingly.
Gyroorbitants were widely used in the 60s-80s, but are now extinct. Simple angular velocity sensors made it possible to effectively measure the rotation of the vehicle, and the on-board computer could easily determine the position of the ship from these data.

Ion sensor

It was a nice idea to supplement the infrared vertical with an ion sensor. In low Earth orbits, there are atmospheric molecules that can be ions - carrying an electrical charge. By installing sensors that record the flow of ions, you can determine which side the ship is flying forward in orbit - there the flow will be maximum:

Scientific equipment for measuring the concentration of positive ions

The ion sensor worked faster - it took almost a whole orbit to build an orientation with a gyroorbitant, and the ion sensor was able to build an orientation in ~10 minutes. Unfortunately, in the area of ​​South America there is a so-called “ion well”, which makes the operation of the ion sensor unstable. According to the law of meanness, it is in the area of ​​South America that our ships need to focus on braking for landing in the Baikonur area. Ion sensors were installed on the first Soyuz, but they were abandoned soon enough, and now they are not used anywhere.

Star sensor

One axis on the Sun is often not enough. For navigation, you may need another bright object, the direction of which, together with the axis to the Sun, will give the desired orientation. The star Canopus became such an object - it is the second brightest in the sky and is located far from the Sun. The first spacecraft to use a star for orientation was Mariner 4, which launched to Mars in 1964. The idea turned out to be successful, although the star sensor drank a lot of the MCC's blood - when constructing the orientation, it was aimed at the wrong stars, and it was necessary to “jump” over the stars for several days. After the sensor finally aimed at Canopus, it began to constantly lose it - debris flying next to the probe would sometimes flash brightly and restart the star search algorithm.
The first star sensors were photocells with a small field of view that could be aimed at only one bright star. Despite their limited capabilities, they were actively used on interplanetary stations. Now technological progress has, in fact, created a new class of devices. Modern star sensors use a matrix of photocells, work in tandem with a computer with a catalog of stars, and determine the orientation of the device based on those stars that are visible in their field of view. Such sensors do not require preliminary construction of a rough orientation by other devices and are able to determine the position of the device regardless of the area of ​​​​the sky to which they are sent.

Typical star trackers

The larger the field of view, the easier it is to navigate

Illustration of the sensor operation - the direction of view is calculated based on the relative positions of the stars according to the catalog data

Advantages:

• Maximum accuracy, can be less than an arc second.


• Does not need other devices, can determine the exact position independently.


• Work in any orbits.

Flaws:

• High price.


• They do not work when the device is rotated quickly.


• Sensitive to light and interference.

Now star sensors are used where it is necessary to know the position of the device very accurately - in telescopes and other scientific satellites.

Magnetometer

A relatively new direction is the construction of orientation according to the Earth's magnetic field. Magnetometers for measuring the magnetic field were often installed on interplanetary stations, but were not used to plot orientation.

The Earth's magnetic field allows you to build orientation along all three axes

"Scientific" magnetometer of the Pioneer-10 and -11 probes

The first digital magnetometer. This model appeared on the Mir station in 1998 and was used in the Philae lander of the Rosetta probe.

Advantages:

• Simplicity, cheapness, reliability, compactness.


• Average accuracy, from arcminutes to several arcseconds.


• You can build orientation along all three axes.

Flaws:

• Subject to interference, including and from spacecraft equipment.


• Does not work above 10,000 km from Earth.

The simplicity and low cost of magnetometers have made them very popular in microsatellites.

Gyro-stabilized platform

Historically, spacecraft often flew unoriented or in solar spin mode. Only in the area of ​​the mission target did they turn on active systems, build orientation along three axes, and complete their task. But what if we need to maintain voluntary orientation for a long time? In this case, we need to “remember” the current position and record our turns and maneuvers. And for this, humanity has not come up with anything better than gyroscopes (measure rotation angles) and accelerometers (measure linear accelerations).
Gyroscopes
The property of a gyroscope to strive to maintain its position in space is widely known:

Initially, gyroscopes were only mechanical. But technological progress has led to the emergence of many other types.
Optical gyroscopes. Optical gyroscopes - laser and fiber optic - are distinguished by very high accuracy and the absence of moving parts. In this case, the Sagnac effect is used - a phase shift of waves in a rotating ring interferometer.

Laser gyroscope

Solid State Wave Gyroscopes. In this case, the precession of a standing wave of a resonating solid is measured. They contain no moving parts and are very accurate.

Vibration gyroscopes. They use the Coriolis effect for operation - vibrations of one part of the gyroscope when turning deflect the sensitive part:

Vibrating gyroscopes are produced in the MEMS version; they are inexpensive and very small in size with relatively good accuracy. It is these gyroscopes that are found in phones, quadcopters and similar equipment. A MEMS gyroscope can also operate in space, and they are installed on microsatellites.

The size and accuracy of gyroscopes is clear:

Accelerometers
Structurally, accelerometers are scales - a fixed load changes its weight under the influence of accelerations, and the sensor converts this weight into an acceleration value. Now accelerometers, in addition to large and expensive versions, have acquired MEMS analogues:

An example of a "large" accelerometer

Micrograph of a MEMS accelerometer

The combination of three accelerometers and three gyroscopes allows you to record rotation and acceleration in all three axes. Such a device is called a gyro-stabilized platform. At the dawn of astronautics, they were only possible on a gimbal and were very complex and expensive.

Apollo gyro-stabilized platform. The blue cylinder in the foreground is a gyroscope. Platform testing video

The pinnacle of mechanical systems were cardless systems, when the platform hung motionless in gas flows. It was high-tech, the result of the work of large teams, very expensive and secret devices.

The sphere in the center is a gyro-stabilized platform. Peacekeeper ICBM guidance system

Well, now the development of electronics has led to the fact that a platform with precision suitable for simple satellites fits in the palm of your hand, it is being developed by students, and even the source code is published.

MARG platforms have become an interesting innovation. In them, data from gyroscopes and accelerometers is supplemented with magnetic sensors, which makes it possible to correct the accumulating error of the gyroscopes. The MARG sensor is probably the most suitable option for microsatellites - it is small, simple, cheap, has no moving parts, consumes little power, and provides three-axis orientation with error correction.
In “serious” systems, star sensors are usually used to correct orientation errors of a gyro-stabilized platform.

One obvious way to improve the efficiency of solar power plants is to use solar tracking systems in them. The development of tracking systems with simple maintenance will significantly improve the technical and economic performance of agricultural facilities and create comfortable working and living conditions for people while ensuring the ecological safety of the environment. Tracking systems can be with one or two axes of rotation of solar panels.

A solar power plant with a tracking system, including a compact photoelectric sun position sensor, consisting of a frame in the shape of a straight triangular prism, on two side faces of which photocells for tracking the sun are located, and on the third face there is a command photocell for turning the modules from west to east. During daylight hours, tracking photocells on the edges of the sensor issue command signals to the control unit for the azimuthal rotation drive of the solar module, which rotates in the direction of the sun using a shaft. The disadvantage of the installation is the insufficient accuracy of tracking the sun.

The solar power plant contains a solar battery with a biaxial orientation system to the sun, on which photoelectric modules containing linear photodetectors located at the foci of cylindrical Fresnel lenses are installed as sun tracking sensors. Signals from photodetectors, using a microprocessor, control the drives of the azimuthal and zenithal orientation system of the solar battery.

The disadvantage of this installation is the insufficient accuracy of tracking the sun, as well as the fact that the tracking sensors occupy part of the active area of ​​the solar battery.

The main goal of the development is to improve the accuracy of the sun tracking sensor for biaxial solar panel orientation systems at any position of the sun in the sky throughout the year.

The above technical result is achieved by the fact that in the proposed sun tracking sensor there is a biaxial solar battery orientation system containing a block of beam-receiving cells installed on a fixed platform, which are made in the form of reverse cones with opaque walls and mounted on the narrow ends of the cones of photovoltaic cells. In this case, the beam-receiving cells are tightly installed on the platform with the formation of a solid angle of 160° and framed by a transparent sphere mounted on the platform, which is installed with an inclination to the horizontal at an angle equal to the geographic latitude of the sensor location.

The tracking sensor is installed on a stationary platform, the normal 6 of which (Fig. 1) is directed to the south. The angle of inclination of the site to the horizontal base corresponds to the geographic latitude of the area next to the solar battery, placed on a mechanical solar orientation system containing zenithal and azimuthal rotation drives using stepper gear motors. The solar battery drives are controlled by a microprocessor that receives electrical impulses from the photoelectric elements of the sensor cells. The microprocessor contains information about the geographic latitude of the location of the solar battery, an electronic clock equipped with a calendar, the signals of which activate the gear motors for the zenithal and azimuthal rotation of the solar battery in accordance with the equation of the movement of the sun in the sky. In this case, the values ​​of the achieved rotation angles of the solar battery based on the signals from the photoelectric elements of the sensor cells are compared with the values ​​​​obtained from the equation of motion of the sun at the current time.

The essence of the sensor design is illustrated in Fig. 1, 2, 3 and 4. In Fig. 1 and 3 show the general diagram of the sensor. In Fig. Figure 2 shows a top view of a transparent sphere and beam-receiving cells. In Fig. Figure 4 shows a diagram of such a cell.

The sun tracking sensor for a biaxial solar panel orientation system contains a platform 1 attached to a horizontal base 5 at an angle a equal to the latitude of the area. A transparent hemisphere 2 with a radius r is attached to the platform 1. In the entire internal space of the sphere 2, beam-receiving cells 3 are closely fixed, having the shape of an inverse cone with opaque walls 7, facing the inner wall of the transparent sphere 2 with a diameter φ, and a diameter d 2 to site 1. The height of cone 3 is equal to the distance h from the inner wall of the sphere 2 to the surface of the platform 1. In the lower part of the cone 3 at a distance of 5d 1 from the upper edge of the cone 3 there is a photoelectric element 4, the electrical signal from which is transmitted to the microprocessor system for controlling the rotation of the solar battery axes (not shown in Fig. 1) . The distance 5d 1 is selected in such a way that the sun's ray 8 is accurately captured on the photoelectric element 4, limited by the opaque walls 7 of the cone 3.

The sun tracking sensor works as follows. The sun's rays 8 penetrate through the transparent sphere 2, the internal space of the cone 3 and fall on the photovoltaic element 4, causing an electric current, which is analyzed by the microprocessor and transmitted to the stepper motor-gear drives of the solar battery orientation system (not shown in the figure). As the sun moves across the sky, its rays 8 gradually turn on the photoelectric elements 3 and contribute to the precise and smooth regulation of the rotation of the solar battery along the azimuthal and zenithal axes.

Laboratory tests of the sensor cell layout using a solar radiation simulator showed acceptable results of cutting off the luminous flux for the accepted values d 1 , d 2 and 5 d x.

The sun tracking sensor of a biaxial solar battery orientation system contains beam-receiving cells made in the form of inverse cones, tightly installed on the site to form a solid angle of 160° and framed by a transparent sphere, allowing for more accurate orientation of solar panels and thereby receiving the greatest amount of electricity from them .

The solar battery rotation system contains a housing, a hollow shaft with a flange for connecting the solar battery, a drive for its rotation, power and telemetric current collectors. The output shaft is functionally divided into a power flange and a shaft with a power current collector. The telemetric current collector is installed on its shaft and connected to the output shaft. The output shaft flange is installed in the housing of the solar battery rotation system on a support bearing with preload or its compression through the support bearing to the housing of the solar battery rotation system by springs. Reliability increases and the weight and dimensions of the device decrease. 1 salary f-ly, 1 ill.

The invention relates to space technology and can be used in the design of a solar array rotation system (SPSB).

The present invention is intended for rotating a solar battery (SB) and transmitting electrical energy from solar batteries to a spacecraft.

A well-known system for rotating solar batteries (SPBS), US patent No. 4076191, consists of a housing, a shaft with two flanges for joining two wings of solar batteries, a drive, and current collectors. Power, transmitting electrical energy, and telemetric, transmitting commands and telemetric information, current collectors are located on the shaft, while the drive turns both wings of the SB. This invention is taken as a prototype.

The disadvantage of this device is the presence of one non-redundant drive and, as a result, reduced survivability of the device. The second disadvantage is the massive design of the shaft, due to the fulfillment of the requirement for the required bending rigidity of the shaft. In addition, a large shaft diameter leads to increased friction and wear of current collectors.

The technical objective of the invention is to increase the reliability of the system, reduce the weight of the structure and increase functionality.

The task is achieved by the fact that in an SPBS having a housing, a drive and a shaft, the output shaft of the device is hollow with a power flange at the end. In this case, the power current collector is located on the output shaft outside, and the telemetric device is installed on its own shaft. The telemetric current collecting device is connected to the output shaft of the SPBS. The output shaft flange is mounted on a support bearing with flat rings or pressed against the housing by springs. The section of the output shaft with the installed power current collector is excluded from the rigid design and has dimensions that are optimal to ensure minimum weight and the required service life of the current collector.

The essence of the invention is illustrated by the drawing, where Fig. 1 shows a general view of the claimed device with a section.

The solar battery rotation system consists of a housing 1, a drive 2, an output shaft 3 mounted on a support bearing 4, a power current collector 6 located on the output shaft 3, and a telemetric current collector 7 mounted on its shaft. The telemetric current-collecting device 7 can be installed in the internal cavity of the output shaft 3 or externally and connected to it. Increased rigidity of the structures is achieved by constantly pressing the shaft 3 to the housing 1 due to the preload of the support bearing or compression by disc springs 8. Increased accuracy of the position of the axis of rotation of the output shaft 3 is achieved by a support bearing with flat support rings 9. The gear wheel 10 is mounted on the shaft 5 of the drive 2. Gear 11 is installed on output shaft 3.

When the SPSB is operating, drive 2 transmits rotation to output shaft 3. Rotation from the drive to output shaft 3 is transmitted by a gear train with gears 10, 11.

Current collectors 6 and 7 transmit electrical energy, commands and signals from the rotating solar array to the spacecraft both when rotating and when stopped. Constant pressure of the output shaft 3 to the housing 1 through the support bearing 4 is ensured by disc springs 8 both during rotation and when the output shaft stops.

Increased survivability of the spacecraft is ensured by the use of one SPSB for each SB wing. Even if the power supply system of one wing fails, the device will receive electrical energy from the other wing and ensure the operation of the main consumers.

The weight reduction of the structure is ensured by the fact that the output shaft 3 is functionally divided into a power flange up to the support bearing 4 and a power current collector shaft. The power flange can be located both inside the SPSB housing and outside, as shown in Fig. 1. The shaft has smaller dimensions, lower weight and increased bending rigidity due to the closure of the power circuit of the structure from the output shaft flange directly to the housing through the support bearing.

The thrust force of the support bearing (or the preload of the support four-point bearing) is selected from the following condition of non-opening of the joint under operating loads:

P>2·K·M/D, where

P - thrust force of the support bearing, Nm;

M - reduced bending moment during normal operation, N;

Reducing the weight of current-collecting devices and increasing their service life is achieved due to the fact that the section of the shaft with the installed power current-collecting device is excluded from the rigid structure and has dimensions that are optimal for the current-collecting device. A capsule-type telemetric current collector device is installed on its shaft, for example, inside the output shaft or is connected externally and has a minimum mass. The increased service life of current collectors is achieved by the possibility of implementing them with a minimum diameter of sliding rings and, accordingly, reduced friction.

Lower friction losses of current collectors make it possible to reduce the drive power, which leads to a reduction in the weight of the drive part of the SPSB.

Currently, the enterprise has released design documentation for the SPSB of the declared design and conducted ground-based experimental testing of the system. Tests have shown a significant reduction in the weight of the system, an increase in service life, an increase in the rigidity characteristics and reliability of the system.

1. A solar battery rotation system having a housing, a hollow shaft with a flange for connecting the solar battery, a drive for its rotation, power and telemetric current collectors, characterized in that the output shaft is functionally divided into a power flange and a shaft with a power current collector, and a telemetric the current collecting device is installed on its shaft and connected to the output shaft, while the output shaft flange is installed in the housing of the solar battery rotation system on a support bearing with preload or its preload through the support bearing to the housing of the solar battery rotation system by springs.

2. The device according to claim 1, characterized in that the preload or preload force of the support bearing is selected from the following condition of non-opening of the joint under operating loads:
P>2·K·M/D,
where P is the preload or preload force of the support bearing, Nm;
K - safety factor for external loads;
M - reduced bending moment during normal operation, N;
D - working diameter of the support bearing (by balls), m.

Similar patents:

The invention relates to equipment of spacecraft (SV) and, in particular, to movable structural elements of the spacecraft that have an electrical connection with the spacecraft control system, for example solar batteries (SB), antennas, movable covers, etc.

The invention relates to control of the orientation of a spacecraft (SV) with solar panels (SB) fixed relative to the spacecraft body. .

The invention relates to the field of space technology and can be used to determine and control the integral parameters of radiant heat transfer of the planet around which the spacecraft (SV) orbits.

The invention relates to space technology and can be used in the design of remote structures of spacecraft, mainly antennas and solar panels. The solar battery strut contains a two-link mechanism, on the common two links of the axis of which a torsion spring with cocking devices is installed. One link is installed on the frame of the solar array, and the other on the body of the spacecraft. A spring-loaded rod is located perpendicular to the axis on one of the links for fixing in the final position. At the end of the spring-loaded rod, a rocker arm is installed with the possibility of rotation, at both ends of which rolling bearings are rigidly fixed, interacting with the conical grooves of the copiers, rigidly mounted on the link opposite the spring-loaded rod. The links of the two-link mechanism have holes for a device for fixing the initial position of the links, secured by means of a threaded connection. EFFECT: increased reliability in the operation of the strut and simplification of the process of installing the solar battery on the spacecraft body. 13 ill.

The invention relates to power supply systems for spacecraft (SC) using solar panels (SB). The method consists in determining a given angle of the SB, measuring its current angle and calculating the calculated angle from the angular velocity of the SB and the time of its rotation. The angles of acceleration (αASG) and braking (αBRAKE) SB are determined. The SB is rotated until the release threshold is reached (αOTP ≈ αTORM), when the mismatch between the specified and calculated SB angles stops. Before starting control, the specified angle is remembered and the initial value of the calculated angle is taken as the reliable value of the current angle. The mismatch threshold (αPR) of these angles is set based on the angles αRAZG and αTORM, as well as the minimum permissible and maximum possible SB currents. The circle of the angle sensor is divided into equal discrete sectors (DS) of size σ under the condition: α ACCELERATION + αBRACK< σ < αПР. Биссектрисы ДС принимают за измеряемые значения. Задают период определения достоверного значения текущего угла на порядок и более превышающим максимальную длительность сбоя информации датчика и менее минимального интервала следования сбоев. Разбивают данный период на четыре равных интервала, и из анализа измеренных и запомненных значений на этих интервалах сбрасывают или формируют сигнал достоверности. В последнем случае вращают СБ до достижения рассогласованием между расчетным и заданным углами значения αОТП и тогда запоминают новое значение заданного угла. Техническим результатом изобретения является повышение живучести и эффективности системы управления ориентацией СБ при кратковременных сбоях информации, поступающей от датчика угла СБ. 4 ил.

The invention relates to power supply systems for a spacecraft (SC) using solar panels (SB). The method includes determining the specified and current angles of orientation of the satellite and the angular velocity (ωSV) of the satellite. The calculated angle is also calculated and, before starting to control the SB, it is assigned the value of the measured angle, which is remembered. Rotate the SB in the direction of decreasing the mismatch between the given and calculated angles. The times and angles of acceleration (tARG, αARG) and deceleration (tBREAK, αBREAK) of the power supply are determined, as well as the maximum permissible angle (αMAX) of the deviation of the power supply, based on the minimum permissible and maximum possible currents of the power supply. At these angles, the response threshold (αCP) is set, when exceeded, the specified mismatch is formed. The latter is not taken into account below the release threshold (αOTP), upon reaching which the rotation of the SB is stopped. The calculated angle of the SB is adjusted within one discrete sector (DS) of the circle of rotation of the SB. The magnitude of the DS depends on the angles αRAZG, αTORM and αCP. Depending on αCP and ωSB, the threshold value of the time for monitoring the continuity of changes in information about the angular position of the SB is set. This monitoring time is counted if the current measured angle differs from the stored angle by more than one DS, and is stopped otherwise. Set the threshold time for controlling the direction of rotation of the SB depending on tRAZG, tBREAK, αMAX, ωSB and the value of the DC. This time is counted at zero continuity control time, if the sign of the discrepancy between the measured and stored angles of the SB does not correspond to the specified direction of rotation of the SB. Otherwise, the countdown is stopped and the rotation direction control time is reset to zero. In this case, at the moment of changing the current measured angle by one DS, the calculated angle is set to the value of the boundary between the DS and the stored angle is assigned a new value of the measured angle. If the continuity control time or the rotation direction control time exceeds its threshold value, then a failure signal is generated and control of the SB is stopped. The technical result of the invention is to increase the survivability and efficiency of the SB attitude control system. 3 ill.

The invention relates to power supply systems for a spacecraft (SC) using solar panels (SB). The method includes determining a given orientation angle of the solar panel to the Sun from the measured angular position of the normal to the working surface of the solar panel and calculating the calculated angle relative to the specified position of the normal. Rotate the SB in the direction of decreasing the mismatch between the given and calculated angles. The angles of acceleration (αASG) and braking (αBRAKE) SB are determined. The calculated angle is adjusted at the moments when the angle sensor values ​​change by the value of the discrete sector (DS) of rotation of the SB. The actuation (αSR) and release thresholds (αOTP) are set, stopping the rotation of the SB if the discrepancy between the given and current angles begins to increase, but not more than αSR. The angular velocity of rotation of the SB is set to be an order of magnitude higher than the maximum angular velocity of the spacecraft's rotation around the Earth, and the value of DS is less than αCP. Set the working angle (αRAB) SB from the condition: αSR< αРАБ < (αГОР - 2·(αРАЗГ + αТОРМ)). Присваивают заданному углу значение углового положения ближайшего к нему луча угла αРАБ, если направление на Солнце в проекции на плоскость вращения указанной нормали находится вне αРАБ. Если угловое положение данной нормали находится вне αРАБ, изменяясь в направлении увеличения угла относительно ближайшего к нему луча угла αРАБ, то формируют сигнал отказа и прекращают управление СБ. Техническим результатом изобретения является исключение заклинивания и поломки панели СБ или бортового оборудования КА, при обеспечении максимально возможного тока в условиях ограничений на углы поворота СБ (напр., от 90° до 180°). 3 ил.

The invention relates to electrical engineering, in particular to devices for generating electrical energy by converting light radiation into electrical energy, and can be used in the creation and production of small-sized spacecraft with solar batteries (SB). The technical result of the invention is: increasing the resistance of the power supply to thermal shocks, to the effects of mechanical and thermomechanical loads, increasing the manufacturability of the design, increasing the active life of spacecraft power supplies, increasing functionality by expanding the temperature range of operation and optimizing the design of the power supply, simplifying the switching system, which is achieved by increasing the strength of the connection of shunt diodes and solar cells, increasing the reproducibility of the manufacturing process of spacecraft solar panels by optimizing the manufacturing technology of shunt diodes and solar cells, as well as switching buses connecting solar cells and shunt diodes, which are made multilayer. A solar battery for small spacecraft contains: panels with modules with solar cells (SCs) glued to them, a shunt diode; switching busbars connecting the front and reverse sides of the shunt diode with the solar cell, while the shunt diode is installed in a cutout in the corner of the solar cell, while the switching busbars are made multilayer, consisting of molybdenum foil, on both sides of which a layer of vanadium or titanium, a layer of nickel and layer of silver respectively. 2 n. and 5 salary f-ly, 4 ill., 3 tables.

The invention relates to controlling the movement of spacecraft (SC) using solar radiation pressure forces distributed over the working areas of the SC. The latter are formed in the form of flat parallel optically transparent droplet flows. The distance between drops of radius R in each flow along it (Sx) and in its frontal-transverse direction (Sy) is a multiple. The number of threads is. By displacing the flows relative to each other in the direction of their movement at a distance, the flows of the droplet sheet are formed in number. Each of these flows is displaced relative to the previous one in the frontal-transverse direction by a distance. This creates opacity in the frontal-transverse direction and transparency in the direction of the plane perpendicular to the flow. The unit distributed force of light pressure is regulated by changing the radius and the number of drops arriving at the point of its application per unit time. The magnitude of the total impact is adjusted by changing the number of drip jets. The technical result of the invention is aimed at increasing the efficiency of using distributed external light pressure forces by reducing their disturbing effect on the relative motion of the spacecraft. 3 ill., 1 tab.

The invention relates to control of the motion of a spacecraft (SV), on which a heat-emitting radiator and a solar battery (SB) are located. The method includes performing a spacecraft flight in orbit around a planet with the solar system turning to a position corresponding to the alignment of the normal to the working surface of the satellite with the direction to the Sun. The orbital orientation of the spacecraft is constructed, in which the SB rotation plane is parallel to the spacecraft orbital plane and the SB is located relative to the orbital plane from the side of the Sun. The altitude of the spacecraft's orbit and the angle between the direction to the Sun and the plane of the spacecraft's orbit are determined. Determine the value (β*) of this angle at which the duration of the shadow part of the turn is equal to the required time for heat release by the radiator on the turn. The orbital orbits at which the current value of a given angle is greater than β* are determined. On these turns, the SB is rotated around the transverse and longitudinal axes of rotation until the conditions for shading the SB radiator are achieved. At the same time, they ensure minimal deviation of the orientation of the working surface of the solar system towards the Sun. The orbital flight of the spacecraft is carried out in a near-circular orbit with an altitude of no more than a certain calculated value. The technical result of the invention is to increase the efficiency of the radiator by creating conditions for its natural cooling when the solar system is shaded in any position of the spacecraft on the orbit. 3 ill.

The invention relates to space technology and can be used in the design of a solar battery rotation system

The invention relates to the power supply of spacecraft (SC) through solar panels (SB), providing useful power from both the working and rear surfaces. The proposed system contains a device for turning the solar panel, an amplifying-converting device, a control unit for the orientation of the solar panel towards the Sun, a block for turning the solar panel to a given position, a block of current regulators, a current sensor, and a control unit for the power supply system. The system additionally includes measurement blocks: the height of the spacecraft orbit, the orientation of the spacecraft and the angle of elevation of the Sun above the Earth's horizon visible from the spacecraft. There is a block for setting the maximum current value generated by the solar system under the influence of direct solar radiation. Blocks have also been introduced to determine: the moments when radiation reflected from the Earth hits the working surface of the solar panel, the moments when radiation reflected from the Earth hits the back surface of the solar panel, the moments when additional electricity is generated by the solar panel under the influence of radiation reflected from the Earth, the angle of rotation of the solar panel and the area of ​​the part of the working surface illuminated by solar radiation SB. The circuit also includes two keys and elements NOT and OR. The technical result of the invention is to increase the solar power output by more completely utilizing solar radiation reflected from the Earth and arriving at the working and rear surfaces of the solar power, taking into account possible shadowing of the solar power surface by spacecraft design elements. 8 ill.

Drawings for RF patent 2341421

The invention relates to the field of space technology, namely to power supply systems (SES) of spacecraft (SC), and can be used to control the position of their solar panels (SB).

To ensure high efficiency of the solar system, most spacecraft are equipped with a system for their automatic orientation to the Sun (see, pp. 190-194; , p. 57). The composition of such a system, taken as an analogue, includes solar sensors, logic converting devices and electric drives that control the position of the solar system. When the system is operating, the solar panel panels are oriented in such a way that the angle between the normal to their illuminated working surface and the direction to the Sun is a minimum value, which ensures the maximum flow of electricity from the solar panel.

The disadvantage of this system for controlling the position of the spacecraft SB is that it does not provide for the operations of placing the SB in fixed design positions, for example, to protect against the negative impact of environmental factors (EFF). FWS can be streams of high-energy particles of solar radiation or streams of gases emerging from operating spacecraft orientation engines.

The closest analogue, adopted as a prototype, is the position control system of the SB spacecraft, described in, p.6.

The block diagram of the system contains a solar panel, on the rigid substrate of the housing of which there is a photovoltaic battery unit (PVB), a solar panel rotating device (UPSB); amplification-converting device (ACD); control unit for SB orientation towards the Sun (BUOSBS); block for turning the SB into a given position (BRSBZP); current regulator block (BRT), AB block (BAB); charger for battery (ZRU AB); unit for generating commands for battery charging (BFKZ AB); load current sensor (LCS); power supply system control unit (BUSES); power supply bus (SE). In this case, the output of the BSE is connected to the input of the BRT. The output of the BRT is connected to the SE. The BAB is connected to the ShE by its input through the AB closed switchgear. The AB switchgear is connected by its first input to the ShE, and the DTN output is connected to the second input of the AB switchgear, the input of which is connected, in turn, to the ShE. The BAB with its output is connected to the first input of the BFKZ AB, and the first output of the BUSES is connected to the second input of the specified block. The output of the BFKZ AB is connected to the third input of the ZRU AB. The second and third outputs of the BUSES are connected, respectively, to the first inputs of the BUSBS and BRSBZP. The UPSB output is connected to the second inputs BUOSBS and BRSBZP. The outputs of BUOSBS and BRSBZP are connected, respectively, to the first and second inputs of the UPU, the output of which, in turn, is connected to the input of UPSB. Moreover, the UPSB is mechanically connected to the SB.

The essence of the actions implemented by this system is as follows. To maximize the flow of electricity from the SB, the SB panels are rotated into a working position corresponding to the alignment of the normal to their illuminated working surface with the plane formed by the axis of rotation of the SB panels and the direction to the Sun. Next, the moment in time of the beginning of the negative impact of the FVS on the working surface of the SB is determined and the SB panels are rotated at a specified angle between the normal to their illuminated working surface and the direction to the Sun until the time of the beginning of the impact of these factors and the SB panels are returned to their working position after the end of the specified impact.

The electricity generated by the BSE is transferred from the SB to the BRT. Next, the electricity from the BRT is supplied to the ShE SES. In the shadow part of the orbit (in the absence of current from the solar system), the battery switchgear, due to the discharge of the battery unit, compensates for the shortage of electricity on board the spacecraft. Along with this, the AB ZRU charges the BAB through the BFKZ AB. At the same time, information from the DTN is used to carry out charge-discharge cycles in the battery switchgear.

Simultaneously with operation in the spacecraft power supply mode, the system solves the problem of controlling the position of the planes of the solar panel panels. Depending on the spacecraft flight program being executed, priority for SB control is given to one of the BUOSBS or BRSBZP blocks.

Upon command from the BUSES, the BUSBS block controls the orientation of the solar system to the Sun. The input information for the solar control algorithm is: the position of the unit direction vector on the Sun relative to the coordinate axes associated with the spacecraft; the position of the SB relative to the spacecraft body, obtained in the form of current measured values ​​of the angle between the current position of the normal to the working surface of the SB and the direction to the Sun from angle sensors (AS) installed on the UPSB. When the SB is oriented towards the Sun, 0. The output information of the control algorithm is commands to rotate the SB relative to the axis of the UPSB output shaft and commands to stop rotation. UPSB remote controls provide discrete signals about the position of the safety system. The discrete size determines the accuracy of the orientation of the satellite.

BRSBZP controls the SB with the help of BUSES according to program settings. The SB control algorithm based on software settings allows you to install the battery in any required position, specified by the required angle value = 2. At the same time, to control the angle of rotation in the BRSBZP, information from the UPSB remote control is also used.

UPU plays the role of an interface between BUOSBS, BRSBZP and UPSB.

It is known (see, p. 272) that solar radiation arriving at the Earth is reflected from its surface, from clouds, and scattered by the atmosphere. The energy of reflected radiation, concentrated in the spectral range of the solar cell sensitivity region, is perceived by the solar cell and increases its output power.

Thus, in the illuminated part of the spacecraft orbit, the SB, in addition to direct solar radiation, receives radiation reflected from the Earth. The method and system adopted as a prototype have a significant drawback - they do not allow increasing the flow of electricity through the additional use of solar radiation reflected from the Earth.

The task facing the proposed system is to increase the flow of electricity from the solar panel through the additional use of solar radiation reflected from the Earth, arriving at the working and rear surfaces of the solar panels, taking into account the possible shading of the surface of the solar panel by the spacecraft design elements.

The technical result is achieved in that the system for controlling the position of the solar panels of the spacecraft, including a solar battery having a positive output power of the rear surface, with a block of photovoltaic batteries installed on it, a device for rotating the solar panels, an amplifying-converting device, a control unit for the orientation of the solar panels according to direction to the Sun, a block for turning solar panels into a given position, a block of current regulators, a current sensor, a control unit for the power supply system, while the output of the photovoltaic battery block is connected to the input of the current regulator block, the output of which is connected to the input of the current sensor, and the outputs of the orientation control blocks solar panels in the direction of the Sun and turning the solar panels to a given position are connected, respectively, to the first and second inputs of the amplification-converting device, the output of which is connected to the input of the device for turning the solar panels, the output of which is connected to the inputs of the control units for the orientation of the solar panels in the direction of The sun and turning the solar panels into a given position, and the device for turning the solar panels is mechanically connected to the solar battery; an additional block for measuring the altitude of the spacecraft orbit, a block for measuring the orientation of the spacecraft, a block for measuring the angle of elevation of the Sun above the Earth's horizon visible from the spacecraft, a task block the maximum value of the current generated by solar panels under the influence of direct solar radiation, a block for determining the moments of radiation reflected from the Earth hitting the working surface of solar panels, a block for determining the moments of radiation reflected from the Earth hitting the back surface of solar panels, a block for determining the moments of generation of additional electricity by solar batteries under influence of radiation reflected from the Earth, a block for determining the angle of rotation of solar panels, a block for determining the area of ​​the part of the working surface of solar panels illuminated by solar radiation, two switches and elements NOT and OR, while the output of the current sensor is connected to the first inputs of the block for determining the angle of rotation of solar panels and the block determining the moments of generation of additional electricity by solar batteries under the influence of radiation reflected from the Earth, the output and the second to fourth inputs of which are connected, respectively, to the input of the NOT element and the outputs of the block for setting the maximum value of the current generated by solar batteries under the influence of direct solar radiation, the OR element and the determination block area of ​​the part of the working surface of the solar panels illuminated by solar radiation, the first and second inputs and output of which are also connected, respectively, to the outputs of the spacecraft orientation measurement unit, the solar panels rotation device and the second input of the solar panels rotation angle determination unit, the output and the third to eighth inputs of which are connected, respectively, to the second input of the block for turning solar panels into a given position and the outputs of the device for turning solar panels, the block for setting the maximum value of the current generated by solar panels under the influence of direct solar radiation, the block for measuring the altitude of the spacecraft orbit, the blocks for determining the moments of radiation reflected from the Earth hitting working and on the back surface of solar panels and a unit for measuring the angle of elevation of the Sun above the Earth's horizon visible from a spacecraft, the output of which is also connected to the first inputs of blocks for determining the moments of radiation reflected from the Earth hitting the working and back surfaces of solar panels, the second inputs of which are connected to the output of the block for measuring the altitude of the spacecraft's orbit, while the outputs of the blocks for determining the moments of radiation reflected from the Earth hitting the working and rear surfaces of the solar panels are also connected, respectively, to different inputs of the OR element, and the output of the power supply system control block is connected to the information inputs of the first and second keys , the control inputs of which are connected to the outputs of the NOT element and the block for determining the moments of generation of additional electricity by solar panels under the influence of radiation reflected from the Earth, respectively, and the outputs of the first and second keys are connected, respectively, to the second input of the block for controlling the orientation of solar panels towards the Sun and the ninth input of the block determining the rotation angle of solar panels.

The proposed invention applies to a class of spacecraft whose solar panels can be shaded by spacecraft structural elements, as well as the solar panels of which have a positive output power when illuminated from the rear surface of the solar panels.

The proposed technical solution achieves an increase in the current generated by SBs that have a positive output power of the rear surface of the SB panels, due to the additional use of solar radiation reflected from the Earth, incident on the working and rear surfaces of the SB panels. To do this, when the spacecraft is in the illuminated part of the orbit, the normal to the working surface on the solar panel is oriented towards the Sun and the time intervals are determined when solar radiation reflected from the Earth arrives either at the working surface or at the rear surface of the solar panels. Then the solar panel is rotated in such a way as to ensure maximum electricity generation from the total illumination of the solar panel by direct solar radiation arriving on the working surface of the solar panels, and radiation reflected from the Earth, arriving on the working or back surface of the solar panels.

The essence of the proposed invention is illustrated in Figs. 1-8, which show: in Figs. 1 and 2 - illumination diagrams for solar panels with direct and reflected solar radiation from the Earth for cases when the radiation reflected from the Earth arrives, respectively, at the working and rear surfaces of solar panels ; Figs. 3 and 4 show SB lighting diagrams in the proposed system; Fig. 5 is a diagram of the geometric construction, explaining the definition of the angle entered below; Fig.6 is a diagram of a geometric construction that explains the determination of the illuminated area of ​​the working surface of the SB, taking into account the shading of the SB; Fig. 7 is a block diagram of the proposed system; Fig. 8 is a graph of the arrival of electricity from the SB of the Russian segment (PC) of the international space station (ISS).

Let us explain the actions implemented by the proposed system.

In Figs. 1-4, which explain the described solar system lighting schemes, all constructions are made in the plane formed by the spacecraft radius vector and the direction to the Sun, and the following symbols are introduced:

N - normal to the working surface of the SB panels;

S, PC, BC * - direction vectors to the Sun;

O - center of the Earth;

OR - radius vector of the spacecraft;

OB - radius of the Earth;

B is the point from which the reflected radiation flow enters the spacecraft;

The angle between the directions from the spacecraft to the Sun and to point B;

MM * - horizon line at point B;

S and are the angle of incidence and angle of reflection from the Earth of solar radiation arriving at the spacecraft;

PD - direction from the spacecraft to the Earth's horizon;

B * - point of contact with the Earth by line PD;

g is the angle of elevation of the Sun above the Earth’s horizon visible from the spacecraft;

Q z is the half-angle of the Earth's disk visible from the spacecraft;

The angle between the directions RO and RV;

Q sb is the half-angle of the sensitivity zone of the working surface of the SB panels, measured from the normal N (indicated only in Figs. 1 and 3);

Angle between N and S (indicated only in Figs. 3 and 4);

In Figs. 2 and 4 it is additionally indicated:

N O - normal to the back surface of the SB panels;

S O - anti-solar direction;

The angle between the direction N o and the direction from the spacecraft to point B;

Q O - half-angle of the sensitivity zone of the rear surface of the SB panels, measured from the normal N o .

We consider the current orientation of the SB, at which the normal to the working surface of the SB N is combined with the direction to the Sun S (at the same time N o is combined with S o).

We use the concept of sensitivity zones of each of the considered surfaces of the solar panel panels - areas determined by the design features of the solar panel elements, when illuminated from the side of which the solar panel is capable of generating electric current. We set the sensitivity zone of each surface of the solar panel panels by the value of the half-angle of the zone, measured from the normal to the considered surface of the solar panel:

Q sb - half-angle of the sensitivity zone of the working surface of the SB panels, Q sb<90°,

Q o - half-angle of the sensitivity zone of the rear surface of the SB panels, Q o<90°.

When illuminating the solar system from outside these areas, the generated current is absent or negligibly small.

The arrival of radiation reflected from the Earth to the spacecraft is possible only in the illuminated part of the orbit, while the location of the reflection point (point B) is determined by the ratio of the angles of incidence s and the reflection of solar radiation from the Earth (see, pp. 39-52;).

After the spacecraft exits the Earth's shadow into the illuminated part of the orbit and before the spacecraft enters the Earth's shadow, the radiation reflected from the Earth hits the working surface of the solar panels (case A, shown in Fig. 1).

This section of the orbit is determined by the conditions:

Taking into account the concept of the SB sensitivity zone, the radiation reflected from the Earth is used by the working surface of the SB panels to generate electricity when the following conditions are met:

then the radiation reflected from the Earth hits the working surface of the solar system and its use for generating electricity is also carried out under the condition

When the spacecraft is in the middle part of the illuminated section of the orbit, the radiation reflected from the Earth affects the back surface of the solar panels (case B, shown in Fig. 2). This section of the orbit is determined by the conditions:

Taking into account the concept of the SB sensitivity zone, the radiation reflected from the Earth is used by the back surface of the SB panels to generate electricity when the following conditions are met:

To determine the angle, you can use different techniques.

From the equality of the sums of the angles that make up the angle ORS, it follows:

In case A, the values ​​of the angles g and are close and the formula can be used:

In case B, the angle is small and the values ​​of the angles and (Q z +g) are close, so you can use the formula:

The half-angle of the Earth's disk visible from the spacecraft Q z is determined from the triangle ORV *:

where it is indicated: R e - radius of the Earth, H o - altitude of the spacecraft orbit.

You can also use a more complex method for determining the angle, one of the possible options of which is the following computational procedure.

In Fig. 5, which explains the definition of angle , it is additionally indicated:

K is the vertex of the right angle of the right triangle ORK.

The angle is determined from the right triangles ORK and OVK:

Substituting expressions (14), (18) into (11) and expressing , we obtain the relation for accurately determining the angle:

The angle is related to the angles , s by the relation obtained from the equality of the angles at the secant PB of parallel lines PC and BC *:

In the case where the nature of the reflection surface allows us to assume equality of the angles of incidence and reflection:

The value satisfying equation (23) is found by iteration using the following procedure.

We denote the solution of this equation relative to as o and denote the function on the right side of (23) as:

At the first iteration, we substitute into function (24) the value equal to 1 - some initial approximation of the desired value o. In case A, it is convenient to take the value of the angle g as an initial approximation; in case B, the value of the sum (Q z +g).

We carry out sequentially for steps i=1, 2, 3,... an iterative process, at each i-th step of which we find i+1 - a new approximation to the desired value o - according to the formula

taking into account the areas of definition of the angle: (2) - in case A and (7) - in case B. Moreover, each new approximation will be closer to the desired value o than the previous one.

We stop the iterative process when the difference between the obtained new approximation i+1 and the previous approximation i is less than the required accuracy of calculations (required accuracy of calculation of the value o):

because in the future, each new approximation will differ from the previous approximation by an amount less than . In this case, the desired value o, to which the sequence of successive approximations i+1, i=1, 2, 3,... converges, also differs from the last obtained approximation by no more than . Thus, the desired value of o, taking into account the required accuracy of calculations, is obtained:

This iterative process quickly converges to the desired solution - for example, for the case of controlling the orientation of the ISS PC, described below as an illustration of the application of this technical proposal, the desired value with an accuracy of 1° is achieved already at the 4th step of the iterative process.

In the absence of solar radiation reflected from the Earth hitting the SB, the current I generated by the SB will be determined by the expression (see, p. 109):

where I is the current current generated by the SB;

I s_max is the current generated by the solar panel when the illuminated working surface of the solar panels is oriented perpendicular to the sun's rays in the absence of radiation reflected from the Earth hitting the surface of the solar panels and in the absence of shading of the working surface of the solar panel by the spacecraft design elements.

We assume that the current generated by the SB is proportional to the surface area of ​​the SB panels onto which the radiation affecting the solar cells of the SB falls. Let's denote:

p s - solar radiation flux density;

S s is the area of ​​the part of the working surface of the solar panels that receives solar radiation;

p o - flux density of radiation reflected from the Earth;

S o is the area of ​​the surface of the SB panels on which the radiation reflected from the Earth arrives.

Let us first consider case A, when the radiation reflected from the Earth arrives at the working surface of the SB (Figs. 1 and 3).

In the proposed system, in this section of the orbit, we deviate the normal to the working surface of the SB N from the direction S in the direction from which the radiation reflected from the Earth arrives at the SB, by the calculated value of the angle between N and S (Fig. 3), ensuring the maximum generation of SB electricity from the total impact of direct solar radiation and radiation reflected from the Earth on the working surface of the solar system. This orientation of the SB is carried out by turning N from S towards the center of the Earth (to the side from which the radiation reflected from the Earth arrives at the SB) by the calculated value of the angle determined as follows.

When N deviates from S in the direction from which radiation reflected from the Earth arrives at the solar panel by an angle , the sum P of the effective values ​​of the fluxes of direct solar radiation and radiation reflected from the Earth arriving at the working surface of solar panels is calculated by the formula (see, page .57):

The formula for calculating the value of the angle that delivers the maximum (29) is obtained by setting the derivative of this expression with respect to zero to zero:

Let us express p o S o from relation (29):

Substituting (33) into (32) we get:

Let's denote:

S s_max - maximum working surface area of ​​SB panels.

Under the influence of the total radiation R, the SBs generate a current current I; under the influence of the radiation flux (p s S s_max), the SBs generate a current equal to I s_max. Wherein

Relationship (34) taking into account (36) takes the form:

Now let's consider case B, when the radiation reflected from the Earth arrives at the back surface of the SB (Figs. 2 and 4).

In the proposed system, in this section of the orbit, we deviate the normal to the rear surface of the SB N o from the direction S o in the direction from which the radiation reflected from the Earth arrives at the SB, to the calculated value of the angle between N o and S o (Fig. 4), providing maximum generation of SB electricity from the total impact of direct solar radiation on the working surface of the SB and on the back surface of the SB - radiation reflected from the Earth. This orientation of the SB is carried out by turning N o from S o towards the center of the Earth (in the direction from which radiation reflected from the Earth arrives at the SB), which is equivalent to turning N from S away from the center of the Earth (or towards the direction of the spacecraft radius vector) , by the calculated value of the angle, determined as follows.

When N o deviates from S o in the direction from which the radiation reflected from the Earth arrives at the SB by an angle , the angle between the direction N o and the direction of the radiation reflected from the Earth arriving at the SB at the source (point B) is determined by the relation:

In this case, the sum P of effective values ​​of radiation fluxes arriving on the working surface of solar panels (direct solar radiation) and the back surface of solar panels (radiation reflected from the Earth) is calculated by the formula:

The formula for calculating the value of the angle that delivers the maximum (40) is obtained by setting the derivative of this expression with respect to zero to zero:

Let us express p o S o from relation (40):

Thus, equations (37) and (46) are obtained for finding the optimal angles of rotation of the SB for cases A and B. The solution of these equations is relatively carried out using the iteration method according to the following procedure.

Let us present equations (37) and (46) in the form, respectively:

Let us denote the functions on the right side of (47) and (48) as:

Let us denote the solution of the equation under consideration as o.

At the first iteration, into function (49) we substitute the value equal to 1 - the initial approximation of the desired value o, for which we take 0° (you can also take the current value of the angle between N and S):

For steps i=1, 2, 3,... we carry out an iterative process, at each i-th step of which we find i+1 - a new approximation to the desired value o - according to the formula:

In this case, each new approximation will be closer to the desired value o than the previous one. We stop the iterative process when the difference between the obtained new approximation i+1 and the previous approximation i is less than the required calculation accuracy:

because in the future, each new approximation will differ from the previous approximation by an amount less than . In this case, the desired value o, to which the sequence of successive approximations i+1, i=1, 2, 3,... converges, also differs from the last obtained approximation by no more than .

Thus, the desired value of o, taking into account the required accuracy of calculations, is obtained:

The radiation reflected from the Earth must be taken into account when the condition is met

when, due to radiation reflected from the Earth hitting the working or back surface of the solar panel, the current value of the current from the solar panel exceeds the maximum possible current value obtained in the absence of radiation reflected from the Earth hitting the solar panel, multiplied by a coefficient taking into account the current possible shading of the working surface of the solar panel elements of the spacecraft design.

The current value of the area S s is calculated as follows. In Fig.6, which explains the necessary geometric constructions, it is indicated:

X sb , Y sb are the coordinate axes of the Cartesian coordinate system associated with the SB, the X sb axis is directed normal to the working surface of the SB.

P 1 P 2 - working surface of the SB;

K 1 K 2 - spacecraft structural element shading the working surface of the SB;

P 1 P p - part of the working surface of the SB, shaded by the element K 1 K 2;

R r R 2 - illuminated part of the working surface of the SB;

P k is the extreme point of the projection of element K 1 K 2 onto the working surface of the SB.

Consider the working surface of a rectangular SB. The coordinates of points P 1 (0; y 1) and P 2 (0; y 2) in the coordinate system associated with the SB are constant, and the value of the entire area of ​​the working surface of the SB S s_max is given by the formula:

where L is the linear size of the SB along the Z axis sb of the Cartesian coordinate system associated with the SB.

Based on measurements of the spacecraft orientation parameters and the position of the satellite relative to the spacecraft, we determine the coordinates of the spacecraft structural elements that shade the working surface of the satellite in the coordinate system associated with the satellite. Let us denote the obtained coordinates of the extreme point of the shading element K 1 K 2 in the coordinate system associated with the SB as K 2 (x k; y k).

Then the coordinates of the point P k are equal to (0; y k), and the coordinate y p of the point P p (0; y p) - the point separating the illuminated and shaded parts of the working surface of the SB - is determined by the formula

The current value of the area S s is calculated by the formula:

The block diagram of the proposed system, presented in Fig. 7, contains the following blocks:

1 - SB; 2 - BSE; 3 - UPSB; 4 - UPU; 5 - BUOSBS; 6 - BRSBZP; 7 - BRT;

8 - DT; 9 - BUSES;

10 - unit for measuring the height of the spacecraft orbit (BIVOKA);

11 - spacecraft orientation measurement unit (BIOKA);

12 - unit for measuring the angle of elevation of the Sun above the Earth's horizon visible from a spacecraft (BIUVSVGZ);

13 - block for setting the maximum current value generated by solar panels under the influence of direct solar radiation (BZMTVSBVPSI);

14 - block for determining the moments of radiation reflected from the Earth hitting the working surface of solar panels (BOMPOSIRPSB);

15 - block for determining the moments of radiation reflected from the Earth hitting the back surface of solar panels (BOMPOSITPSB);

16 - block for determining the moments of generation of additional electricity by solar batteries under the influence of radiation reflected from the Earth (BOMGSBDEVOZI);

17 - block for determining the angle of rotation of solar panels (BOUPSB);

18 - block for determining the area of ​​the part of the working surface of solar panels illuminated by solar radiation (BOPOSIRPSB);

19, 20 - first and second keys;

21 - element NOT;

22 - OR element,

in this case, the output of the BSE (2) is connected to the input of the BRT (7). The BRT output (7) is connected to the DT input (8). The outputs of BUOSBS (5) and BRSBZP (6) are connected, respectively, to the first and second inputs of the UPU (4). The output of the UPU (4) is connected to the input of the UPS (3). The output of UPSB (3) is connected to the first inputs of BUOSBS (5) and BRSBZP (6). The DT output (8) is connected to the first inputs BOUPSB (17) and BOMGSBDEVOSI (16). The output and the second to fourth inputs BOMGSBDEVOSI (16) are connected, respectively, to the input of the NOT element (21) and to the outputs of BZMTVSBVPSI (13), the OR element (22) and BOPOSIRPSB (18). The first and second inputs and output of BOPOSIRPSB (18) are also connected, respectively, to the outputs of BIOKA (11), UPSB (3) and the second input of BOUPSB (17). The output and the third to eighth inputs of BOUPSB (17) are connected, respectively, to the second input of BRSBZP (6) and the outputs of UPSB (3), BZMTVSBVPSI (13), BIVOKA (10), BOMPOSIRPSB (14), BOMPOSITPSB (15), BIUVSVGZ (12 ). The output of BIUVSVGZ (12) is also connected to the first inputs of BOMPOSIRPSB (14) and BOMPOSITPSB (15). The second inputs BOMPOSIRPSB (14) and BOMPOSITPSB (15) are connected to the output of BIVOKA (10). The outputs BOMPOSIRPSB (14) and BOMPOSITPSB (15) are also connected, respectively, to different inputs of the OR element (22). The output of the BUSES (9) is connected to the information inputs of the first and second keys (19) and (20). The control inputs of the first and second keys (19) and (20) are connected to the outputs of the element NOT (21) and BOMGSBDEVOSI (16), respectively. The outputs of the first and second keys (19) and (20) are connected, respectively, to the second input of the BUOSBS (5) and the ninth input of the BUOSSB (17).

Figure 7 also shows with a dotted line the mechanical connection of the UPSB (3) with the SB housing (1) through the output shaft of the SB drive.

The system works as follows.

Electricity from the BSE (2) is supplied to the BRT (7), then from which it is supplied to the SE SES of the spacecraft. In this case, the BRT (7) is connected to the DT (8), which measures the current value of the current generated by the SB.

In BIVOKA (10) the value of the spacecraft orbit altitude is measured.

In BIOKA (11) the spacecraft orientation parameters are measured.

BIUVSVGZ (12) determines the value of the angle of elevation of the Sun above the Earth's horizon visible from the spacecraft.

BOMPOSIRPSB (14) determines the moments of time at which radiation reflected from the Earth can hit the working surface of the SB panels. To do this, the fulfillment of condition (5) is checked. This block can also implement a more complex computational scheme, including calculating the angle using formula (12) or using the computational procedure (23)-(27) and checking condition (3). When conditions (5), (3) are met, the BOMPOSIRPSB block (14) generates a command that arrives at the first input of the OR element (22).

BOMPOSITPSB (15) determines the moments of time at which radiation reflected from the Earth can hit the back surface of the SB panels. To do this, the fulfillment of condition (6) is checked. This block can also implement a more complex computational scheme, including calculating the angle using formula (13) or using the computational procedure (23)-(27) and checking condition (10). When conditions (6), (10) are met, the BOMPOSITPSB block (15) generates a command that arrives at the second input of the OR element (22).

When a command is received at any of the two inputs of the OR element (22), a command is generated at the output of the OR element (22) and sent to the corresponding input of BOMGSBDEVOSI (16). Note that the BOMPOSIRPSB (14) and BOMPOSITPSB (15) blocks cannot simultaneously generate commands, because they check the fulfillment of mutually exclusive geometric conditions.

In BOPOSIRPSB (18), the area of ​​that part of the working surface of the solar system that is currently illuminated by direct solar radiation is determined. Based on the input information about the spacecraft orientation parameters coming from the BIOKA (11) and the position angle of the SB relative to the spacecraft coming from the UPSB (3), the BOPOSIRPSB block (18) implements the computational procedure (56)-(57).

In BOMGSBDEVOSI (16) the moments of use of SB radiation reflected from the Earth are determined - the moments of generation of additional electricity by SB under the influence of radiation reflected from the Earth. These moments correspond to the simultaneous fulfillment of condition (54) and the conditions for the radiation reflected from the Earth to hit the working or rear surface of the SB panels (the latter conditions are fulfilled in the BOMPOSIRPSB (14) and BOMPOSITPSB (15) blocks). When condition (54) is simultaneously met and a signal is received from the OR element (22), the BOMGSBDEVOSI block (16) generates a command that arrives at the input of the NOT element (21) and the control input of the key (20).

If condition (54) is not met or a signal is not received from the OR element (22) at the output of BOMGSBDEVOSI (16), the command is not generated. Then the NOT element (21) generates a command sent to the control input of the key (19). In this case, the key (20) is closed, and the key (19) is open.

In this state of the keys (19) and (20), the control command from the BUSES (9) through the open key (19) enters the BUSBS unit (5), which controls the orientation of the SB (1) to the Sun. BUOSBS (5) can be implemented on the basis of the motion and navigation control system (VCS) of the spacecraft (see). The input information for the satellite control algorithm is: the position of the unit direction vector to the Sun relative to the coordinate axes associated with the spacecraft, determined by the algorithms of the kinematic contour of the vessel; the position of the SB relative to the spacecraft body, obtained in the form of the current measured values ​​of the angle with the UPSB remote control (3). The output information of the control algorithm is commands to rotate the SB relative to the axis of the output shaft of the UPSB (3), commands to stop rotation. The UPSB remote control (3) provides signals about the position of the SB (1).

When BOMGSBDEVOSI (16) issues a command arriving at the control input of the key (20) and the NOT element (21), then the NOT element (21) does not generate a command at the control input of the key (19). In this case, the key (20) is open, and the key (19) is closed.

In this state of the keys (19) and (20), the control command from the BUSES (9) through the public key (20) is sent to the BUPSB (17).

When a command is received from BUSES (9) to the input BOUPSB (17), the BOUPSB block (17), depending on the commands received from the blocks BOMPOSIRPSB (14) and BOMPOSITPSB (15), calculates the rotation angle SB = o using computational procedures (47)- (53). In this case, the angle is also calculated using formulas (12), (13) or (19), (23)-(27). For calculations, the values ​​, I, I s_max , S s ​​, g, H o are used, coming from UPSB (3), DT (8), BZMTVSBVPSI (13), BOPOSIRPSB (18), BIUVSVGZ (12), BIVOKA (10). =

The implementation of the blocks BOMPOSIRPSB (14), BOMPOSITPSB (15), BOMGSBDEVOSI (16), BOUPSB (17), BOPOSIRPSB (18) is possible both on the basis of hardware and software of the spacecraft flight control center (MCC) and on board the spacecraft. An example of the implementation of BUSES (9) can be the radio means of the service control channel (SCU) onboard systems of the Yamal-100 spacecraft, consisting of an earth station (ES) and on-board equipment (BA) (see description in). In particular, the BA SKU together with the 3D SKU solves the problem of issuing digital information (DI) to the on-board digital computer system (OBDS) of the spacecraft and its subsequent acknowledgment. BCWS, in turn, controls the blocks BUOSBS (5), BOUPSB (17), BRSBZP (6).

UPU (4) plays the role of an interface between BUOSBS (5), BRSBZP (6) and UPSB (3) and serves to convert digital signals into analogue ones and amplify the latter.

BIVOKA (10), BIOKA (11), BIUVSVGZ (12) can be made on the basis of sensors and equipment of the spacecraft (see,). The implementation of BZMTVSBVPSI (13), BOMPOSIRPSB (14), BOMPOSITPSB (15), BOMGSBDEVOSI (16), BOUPSB (17), BOPOSIRPSB (18) can be carried out on the basis of BTsVS. Keys (19), (20), NOT element (21), OR element (22) can be made in the form of elementary analog circuits. SB (1), BFB (2), UPSB (3), UPU (4), BUOSBS (5), BRSBZP (6), BRT (7), DT (8) can be made on the basis of SES elements (see) .

Thus, an example of the implementation of the fundamental blocks of the system is considered, based on the results of which a decision is made and the proposed operations are implemented.

Let us describe the technical effect of the proposed inventions.

The proposed technical solutions ensure maximum generation of electricity from the total impact on the solar panel of direct solar radiation arriving on the working surface of the solar panels, and radiation reflected from the Earth arriving at the working or rear surface of the solar panels, taking into account possible shadowing of the working surface of the solar panel by the spacecraft design elements. In this case, an increase in the receipt of electricity from the solar panel is achieved by increasing the use of the working and back sides of the surfaces of solar panels of radiation reflected from the Earth by performing, at the proposed time intervals, the proposed turns of the solar panel from the direction of the Sun in a given direction, determined by the direction of arrival of radiation reflected from the Earth to the spacecraft, to the calculated angle determined by the proposed method.

For illustration, Fig. 8 shows a graph of the arrival of electricity from the SB PC MKC I(A) versus time t (s) during an orbital revolution while maintaining the orientation of the SB to the Sun: 02.02.2004, orbit 1704, time 17.35-19.06 DVM, orientation of the ISK (see). The graph shows the current level I s_max and marks the time intervals T 1, T 2 located at the beginning and end of the illuminated part of the orbital revolution and corresponding to the moments when condition (3) is met, and the time interval T o located in the middle part of the illuminated part of the orbit and corresponding to the moments of fulfillment of condition (10). The graph illustrates that condition (54) is satisfied at these intervals, i.e. On the surface of the SB panels, radiation reflected from the Earth is additionally received, and rotating the SB by the calculated angle = o allows increasing the generation of SB electricity under the influence of the total radiation arriving on the surface of the SB panels.

LITERATURE

1. Eliseev A.S. Space flight technology. M.: Mechanical Engineering, 1983.

2. Rauschenbach G. Handbook for the design of solar panels. M.: Energoatomizdat, 1983.

3. Kovtun V.S., Solovyov S.V., Zaikin S.V., Gorodetsky A.A. A method for controlling the position of solar panels of a spacecraft and a system for its implementation. Description of the invention for RF patent No. 2242408 according to application 2003108114/11 dated March 24, 2003.

4. Kroshkin M.G. Physical and technical foundations of space research. - M.: Mechanical Engineering. 1969.

5. Kondratyev K.Ya. Actinometry. - M.: Gidrometeoizdat. 1965.

6. Grilikhes V.A., Orlov P.P., Popov L.B. Solar energy and space flights. M.: Nauka, 1984.

7. Spacecraft motion control and navigation system. Technical description. 300GK.12Yu. 0000-ATO. RSC Energia, 1998.

8. Earth station of the service control channel of the Yamal spacecraft. Manual. ZSKUGK.0000-0RE. RSC Energia, 2001.

9. Onboard equipment of the service control channel of the Yamal spacecraft. Technical description. 300GK.15Yu. 0000A201-OTO. RSC Energia, 2002.

10. Engineering reference book on space technology. Publishing house of the Ministry of Defense of the SSR, M., 1969.

11. Spacecraft power supply system. Technical description. 300GK.20Yu. 0000-ATO. RSC Energia, 1998.

12. Rulev D.N., Stazhkov V.M., Korneev A.P., Panteleimonov V.N., Melnik I.V. Assessment of the efficiency of solar panels of the Russian segment of the international space station based on telemetric information // Proceedings of the XXXIX Readings dedicated to the development of the scientific heritage and the development of ideas of K.E. Tsiolkovsky (Kaluga, September 14-16, 2004). Section “Problems of rocket and space technology”. - Kazan: Kazan State University named after. V.I.Ulyanov-Lenin. 2005.

CLAIM

A system for controlling the position of solar panels of a spacecraft having blocks of photovoltaic batteries installed on them with a positive output power of the rear surface, containing a device for rotating solar panels, an amplifying-converting device, a control unit for the orientation of solar panels towards the Sun, a unit for turning solar panels into a given position , a current regulator block, a current sensor, a power supply system control unit, wherein the output of the photovoltaic battery unit is connected to the input of the current regulator unit, the output of which is connected to the input of the current sensor, and the outputs of the control units for the orientation of solar panels towards the Sun and the rotation of solar panels in the specified position are connected, respectively, to the first and second inputs of the amplification-converting device, the output of which is connected to the input of the solar panel rotation device, the output of which is connected to the inputs of the control units for the orientation of the solar panels towards the Sun and the rotation of the solar panels to a given position, and the solar panel rotation device batteries is mechanically connected to the specified solar battery, characterized in that it additionally includes a unit for measuring the altitude of the spacecraft orbit, a unit for measuring the orientation of the spacecraft, a unit for measuring the angle of elevation of the Sun above the Earth's horizon visible from the spacecraft, a unit for setting the maximum value of the current generated by the solar batteries under the influence of direct solar radiation, a unit for determining the moments when radiation reflected from the Earth hits the working surface of solar batteries, a unit for determining when radiation reflected from the Earth hits the back surface of solar batteries, a unit for determining when solar batteries generate additional electricity under the influence of radiation reflected from the Earth, a block for determining the angle of rotation of solar panels, a block for determining the area of ​​the working surface of solar panels illuminated by solar radiation, two keys and elements “NOT” and “OR”, while the output of the current sensor is connected to the first inputs of the block for determining the angle of rotation of solar panels and the block for determining the moments of generation solar batteries of additional electricity under the influence of radiation reflected from the Earth, the output and inputs - from the second to the fourth - of which are connected, respectively, to the input of the "NOT" element and the outputs of the block for setting the maximum value of the current generated by the solar batteries under the influence of direct solar radiation, the "OR" element and a block for determining the area of ​​the part of the working surface of solar panels illuminated by solar radiation, the first and second inputs and output of which are also connected, respectively, to the outputs of the unit for measuring the orientation of the spacecraft, the device for rotating the solar panels and the second input of the block for determining the angle of rotation of the solar panels, the output and inputs - with third through eighth - which are connected, respectively, to the second input of the block for turning the solar panels into a given position and the outputs of the device for turning the solar panels, the block for setting the maximum value of the current generated by the solar panels under the influence of direct solar radiation, the block for measuring the altitude of the spacecraft orbit, the blocks for determining the moment of impact radiation reflected from the Earth onto the working and back surfaces of the solar panels and a unit for measuring the angle of elevation of the Sun above the Earth's horizon visible from the spacecraft, the output of which is also connected to the first inputs of the blocks for determining the moments when radiation reflected from the Earth hits the working and back surfaces of the solar panels, the second inputs of which are connected to the output of the unit for measuring the height of the orbit of the spacecraft, while the outputs of the units for determining the moments when radiation reflected from the Earth hits the working and rear surfaces of the solar panels are also connected, respectively, to different inputs of the "OR" element, and the output of the power supply system control unit is connected with information inputs of the first and second keys, the control inputs of which are connected to the outputs of the "NOT" element and the unit for determining the moments of generation of additional electricity by solar panels under the influence of radiation reflected from the Earth, and the outputs of the first and second keys are connected, respectively, to the second input of the solar orientation control unit batteries in the direction of the Sun and the ninth input of the block for determining the angle of rotation of solar batteries.

Owners of patent RU 2322373:

The inventions relate to power supply of spacecraft (SC) using solar panels (SB). The proposed method involves rotating the solar panels into a working position corresponding to the alignment of the normal to their illuminated surface with the plane formed by the axis of rotation of the solar panels and the direction to the Sun. At the same time, the flux densities of solar electromagnetic radiation and high-energy particles are measured, determining the moments of the onset of solar activity and the arrival of these particles on the surface of the spacecraft. Additionally, the moments of appearance of precursors of the negative impact of the flows of these particles on the spacecraft are determined. At these moments, the spacecraft's onboard batteries are charged to the maximum level. When particle flux densities exceed threshold values, solar panel panels are deployed to an angle between the specified normal and the direction to the Sun, corresponding to the minimum area of ​​influence of particle fluxes on the surface of solar panels. The shortage of electricity on board the spacecraft is covered by discharging the batteries. When the minimum permissible charge level of these batteries is reached, they are disconnected from the load. After the impact of particles on the spacecraft has ended, the SB panels are returned to their working position. The proposed control system includes the necessary blocks and connections between them to perform the operations described above. Moreover, it includes a block for determining the required current from the solar system, a block for determining the moments of the appearance of harbingers of the negative impact of high-energy particles on the spacecraft, and a block for setting the permissible charge level of the batteries. The technical result of the inventions is to weaken the negative impact of high-energy particle flows on the working surface of the solar panel by maximizing the angle of the “protective” turn of the solar panel from the direction of these flows from the Sun. 2 n.p. f-ly, 1 ill.

The invention relates to the field of space technology, namely to power supply systems (SES) of spacecraft (SC), and can be used to control the position of their solar panels (SB).

There is a known method for controlling the position of SB panels, adopted as an analogue (see, pp. 190-194). The essence of the method is as follows. The SB panels are oriented in such a way that the angle between the normal to their illuminated working surface and the direction to the Sun is a minimum value, which ensures the maximum flow of electricity from the SB.

To ensure high efficiency of the solar system, most spacecraft are equipped with a system for their automatic orientation to the Sun. Such a system includes solar sensors, logic converting devices and electric drives that control the position of the solar system.

The disadvantage of this method and the spacecraft SB position control system is that their actions do not provide protection from the negative impact of environmental factors (EFF) on the working surfaces of the SB panels, such as, for example, protection from gases escaping from operating jet engines (RE). ) spacecraft (see, p. 311-312; , p. 2-27), and fluxes of protons and electrons of high energies of cosmic rays of solar electromagnetic radiation (EMR) during periods of high solar activity (see, p. 323; , p. .31, 33).

The closest analogue, adopted as a prototype, is the method of controlling the position of the satellite satellite, described in. The essence of the method is as follows.

The SB panels are rotated into a working position that ensures the spacecraft is supplied with electricity, corresponding to the alignment of the normal to its illuminated working surface with the plane formed by the axis of rotation of the SB panels and the direction to the Sun. Next, the moment in time of the beginning of the negative impact of the FVS on the working surface of the SB is determined and the SB panels are rotated until the time when the impact of the specified factors begins and the SB panels are returned to their working position after the end of the specified impact. To do this, the density of the current flux of solar electromagnetic radiation is measured and, based on the measured values, the moment in time of the beginning of solar activity is determined, and the moment in time when particles reach high energy levels on the surface of the spacecraft is determined. At a specified point in time, the flux density of high-energy particles - protons and electrons - is measured and the measured values ​​are compared with threshold values. If the measured values ​​exceed the threshold values ​​of proton and electron fluxes, the solar panel panels are rotated at the angle between the normal to their illuminated working surface and the direction to the Sun α s_min, corresponding to the minimum area of ​​influence of high-energy particle fluxes on the solar panel surface, determined by the relation:

α s min =arccos(I n /I m),

where I n - load current from spacecraft consumers;

I m - maximum current generated when the illuminated working surface of the solar panels is oriented perpendicular to the sun's rays,

in this case, the moment in time when the measured values ​​exceed the upper threshold value of the flux density of the specified high-energy particles is taken as the moment of time when the SB panels begin to turn, and the moment in time when the flux density of high-energy particles becomes lower than the upper threshold is taken as the moment of time when the SB panels begin to return to their working position threshold value.

SBs in the ISS SES system are the main sources of electricity and ensure the operation of its onboard consumers, including recharging batteries (AB), which are secondary sources of electricity on board the ISS (see). By turning the SB, the area of ​​damage to the working surfaces of the SB by the FVS flow is reduced. It is not possible to completely deploy the SB panels along the damaging FWS flow, because it is necessary to provide the spacecraft and its batteries with electricity generated by the solar power system, - based on this, the area affected by the solar power panels by the flow of high-energy particles is reduced to a minimum by turning the solar power system at an angle α s min, necessary and sufficient to provide on-board consumers with energy.

Based on the necessary sufficiency, for the operation of on-board systems of the spacecraft, the load from consumers I n should not exceed the current current I. Since the current current I from the SB is determined by the expression (see, p. 109)

where I m is the maximum current generated when the illuminated working surface of solar panels is oriented perpendicular to the sun's rays;

α is the current angle between the normal to the working surface of the solar system and the direction to the Sun,

then the current angle α should not exceed the value α s min, calculated by the formula:

The SB position control system for implementing this method, adopted as a prototype, is described in and contains a SB, on the rigid substrate of which there are four photovoltaic batteries (BF 1, BF 2, BF 3, BF 4), a SB rotation device (UPSB); amplification-converting device (ACD); control unit for SB orientation towards the Sun (BUOSBS); block for turning the SB into a given position (BRSBZP); two current regulators (PT 1, PT 2), AB unit (BAB); charger for battery (ZRU AB); unit for generating commands for battery charging (BFKZ AB); load current sensor (LCS); power supply system control unit (BUSES); power supply bus (SE); unit for measuring the density of the current solar EMR flux (BIPEMI); solar activity detection unit (BOSA); block for determining the moment of impact of particles on the spacecraft (BOMVVCH); unit for measuring the density of high energy particle fluxes (HIPPCHVE); block for determining the moment of start of SB control based on load currents (BOMVUSBTNZ); SB control unit for load currents (BUSBTNZ). In this case, the SB, through its first output, combining the outputs of BF 1 and BF 4, is connected to the first input of the UPSB, and through the second output, combining the outputs of BF 2 and BF 3, is connected to the second input of the UPSB. The outputs of BUOSBS and BRSBZP are connected, respectively, to the first and second inputs of the UPU, the output of which, in turn, is connected to the third input of UPSB. The first and second outputs of the UPSB are connected respectively to the inputs PT 1 and PT 2, and the outputs PT 1 and PT 2 are connected to the SE. The BAB is connected to the ShE by its input through the AB closed switchgear. In this case, the AB switchgear is connected with its first input to the specified bus, and the accident output is connected to the second input of the AB switchgear, the input of which is connected, in turn, to the ShE. The BAB with its output is connected to the first input of the BFKZ AB, and the first output of the BUSES is connected to the second input of the specified block. The output of the BFKZ AB is connected to the third input of the ZRU AB. The second and third outputs of the BUSES are connected, respectively, to the first inputs of the BUOSBS and BRSBZP. The third output of the UPSB is connected to the second inputs of the BUOSBS and BRSBZP. The BIPEMI output is connected to the BOSA input, the first output of which, in turn, is connected to the BOMVVCH input. The outputs of BOMVVCH and BIPPChVE are connected to the first and second inputs of the BOMVUSBTNZ block, respectively, and the input of BIPPCHVE is connected to the second output of BOSA. The output of BOMVUSBTNZ is connected to the input of BUSES. BUSES with its fourth output is connected to the first input of BUSBTNZ, and the second output of DTN is connected to the second input of BUSBTNZ. The output of BUSBTNZ is connected to the third input of the UPU. In addition, the third output of the UPSB is connected to the third input of the BUSBTNZ.

In the spacecraft power supply mode, the system operates as follows.

UPSB serves for transit transmission of electricity from SB to PT 1 and PT 2. Voltage stabilization on the SES power supply bus is carried out by one of the RTs. At the same time, the other RT is in a state with power transistors closed. In this case, SB generators operate in short circuit mode. When the load power becomes greater than the connection power of the solar power generators, another RT switches to the voltage stabilization mode, and the energy of the unused generators is supplied to the power supply bus of the solar power plant. In certain periods, when the load power may exceed the power of the battery, the battery switchgear, due to the discharge of the battery unit, compensates for the shortage of electricity on board the spacecraft. For these purposes, the battery discharge regulator serves as a battery discharge regulator.

In addition to the specified regulator, the battery charger also contains a battery charge regulator. The charge regulator limits the charging current of the battery at the level of (I cl ±1)A, where I cl is the rated charge current, in case of excess power of the battery and stabilizes the voltage on the SES bus by regulating the charging current of the battery when the power of the battery is insufficient to provide power to the battery charging current (I nc ±1)A. To carry out the specified charge-discharge cycles in the battery switchgear, information from the DTN is used. At the same time, the DVT is connected to the SES in such a way that it measures the load current not only from on-board consumers, but also takes into account the battery charging current. The charge of the BAB is carried out by the ZRU AB through the BFKZ AB.

Simultaneously with operation in the spacecraft power supply mode, the system solves the problem of controlling the position of the planes of the solar panel panels.

Upon command from the BUSES, the BUSBS block controls the orientation of the solar system to the Sun. BUOSBS can be implemented on the basis of the motion and navigation control system (VCS) of the spacecraft (see). In this case, the input information for the satellite control algorithm is: the position of the unit direction vector to the Sun relative to the coordinate axes associated with the spacecraft, determined by the algorithms of the kinematic contour of the vessel; the position of the SB relative to the spacecraft body, obtained in the form of current measured values ​​of the angle α from angle sensors (AS) installed on the UPSB. In this case, the value of α is always measured from the current normal to the working surface of the SB (i.e., when the SB is oriented towards the Sun, α is minimal). The output information of the control algorithm is commands to rotate the SB relative to the axis of the output shaft of the UPSB and commands to stop rotation. UPSB remote controls provide discrete signals about the position of the safety system. The discrete size determines the accuracy of the orientation of the satellite.

In the normal mode of spacecraft orientation, when the direction of the Sun's motion relative to the connected axes of the spacecraft is unchanged, the SB is set relative to the direction to the Sun with an advance in the direction of the Sun's motion by an angle corresponding to several discretes of the remote control. Then the battery remains in this position until the Sun, due to the motion of the spacecraft in orbit, “moves forward” relative to the SB at the appropriate angle. After this, the rotation cycle resumes.

BRSBZP controls the SB with the help of BUSES according to program settings. The SB control algorithm based on software settings allows you to install the battery in any specified position. To do this, a signal is initially issued to the BUOSBS about setting the SB to its original position. Next, using the BUSBZP, the required turn through the angle α z is carried out. At the same time, to control the angle of rotation in the BRSBZP, information from the UPSB remote control is also used.

UPU plays the role of an interface between BUOSBS, BRSBZP, BUSBTNZ and UPSB.

BIPEMI continuously measures current fluxes of solar electromagnetic radiation (EMR) according to the solar activity index F10.7 and transmits them to BOSA. In BOSA, by comparing current values ​​with specified threshold values, the onset of solar activity is determined. According to the command coming from the first output of the BOSA to the input of the BOMVHF, in the indicated last block the moment in time of the possible beginning of the impact of high-energy particles on the spacecraft is determined. From the second output of the BOSA through the input of the BIPPCHVE, a command is issued to begin measuring the flux density of high-energy particles. Information about the moment in time of the possible start of the impact of particles on the spacecraft is transmitted from the output of the BOMVVCH to the BOMVUSBTNZ through its first input. The measured value of the flux density of high-energy particles from the BIPPCHVE is transmitted to the second input of the BOMVUSBTNZ.

In BOMVUSBTNZ, the actual assessment of the negative impact of FVS is carried out by comparing the current measured value of the impact characteristic with threshold values, starting from the time point determined by BOMVUSBTNZ. A necessary condition for receiving a command at the BOMVUSBTNZ output is the presence of two signals - from the BOMVVCH and BIPPCHVE outputs. At the output of BOMVUSBTNZ, the command “start control of the power supply based on load currents” is generated, which is sent to the BUSES.

When BOMVUSBTNZ issues a command to BUSES, the command received from BOMVUSBTNZ has a higher priority than the commands to activate BUOSBS and BRSBZP. Therefore, having received the specified command, BUSES disconnects lower priority blocks from UPSB control and connects BUSBTNZ.

After the command from BOMVUSBTNZ is reset to zero at the BUSES input, the latter rebuilds the logic of its operation. Depending on the spacecraft flight program being executed, priority for SB control is given to one of the BUOSBS or BRSBZP blocks.

BUSBTNZ determines the angle α s_min using expression (2). To calculate the specified angle, the measured values ​​of I n obtained from the DTN are used. In addition, from the UPSB remote control, the specified block receives information about the current value of the SB rotation angle α. Having determined the value of the angle α s_min, the algorithm embedded in BUSBTNZ compares it with the current value of the angle α, calculates the mismatch angle between α and α s_min and the required number of control pulses to activate the control drive SB. Control pulses are transmitted to the control unit. After converting and amplifying the indicated pulses in the UPU, they enter the input of the UPS and set the drive in motion.

The method and system for its implementation, adopted as a prototype, have a significant drawback - they do not provide complete protection of the solar panel surface from the negative effects of high-energy particle flows and, at the same time, do not allow the use of additional opportunities to reduce this negative impact by performing special operations for the preparation of solar panels Spacecraft to operate under conditions of the negative impact of high-energy particle flows on the spacecraft.

The challenge facing the proposed method and system for its implementation is to reduce the negative impact of high-energy particle flows on the SB surface. To do this, by performing special preparatory operations in the spacecraft SES and controlling the SB, it is intended to reduce the area of ​​the SB, which is negatively affected by the flow of these particles.

The technical result is achieved by the fact that in the method of controlling the position of the solar panels of a spacecraft, including turning the solar panels into a working position, ensuring the supply of the spacecraft with electricity corresponding to the alignment of the normal to its illuminated working surface with the plane formed by the axis of rotation of the solar panels and the direction to The sun, measuring the density of the current flux of solar electromagnetic radiation, determining the moment in time when solar activity begins, determining the moment in time when high-energy particles reach the surface of the spacecraft, measuring the flux density of high-energy particles, comparing the measured values ​​of the flux density of high-energy particles with threshold values, turning solar panels batteries by the angle between the normal to their illuminated working surface and the direction to the Sun, corresponding to the minimum area of ​​influence of high-energy particle fluxes on the surface of solar panels while simultaneously providing the spacecraft with electricity, at the moment when the measured values ​​of the high-energy particle flux density exceed threshold values ​​and the return of the panels solar panels into the operating position at the time at which the density of high-energy particle fluxes becomes below threshold values, additionally determine the times of appearance of the precursors of the negative impact of high-energy particle fluxes on the spacecraft, at the time of the appearance of the precursors of the negative impact of high-energy particle fluxes on the spacecraft The device charges the batteries of the spacecraft power supply system to the maximum charge level; if the measured values ​​of the flux density of high-energy particles exceed the threshold values ​​compared with them, the solar panels are rotated until the angle between the normal to their illuminated working surface and the direction to the Sun is reached α s_min_AB, corresponding to the minimum area of ​​influence of high-energy particle flows on the surface of solar panels while simultaneously providing the spacecraft with electricity from solar and rechargeable batteries of the power supply system, determined by the relation:

α s_min_AB =arccos(max(0,I n -I AB )/I m),

where I n is the load current from the spacecraft consumers,

I m - maximum current generated when the illuminated working surface of solar panels is oriented perpendicular to the sun's rays,

I AB - current permissible discharge current of batteries,

and the resulting shortage of electricity on board the spacecraft is compensated by discharging the batteries, while monitoring the charge level of the batteries and, upon reaching the minimum permissible value of the charge level of the batteries, the current value of the permissible discharge current of the batteries is reset and the batteries are disconnected from the external load.

In addition, the problem is solved by the fact that in the system for controlling the position of the solar panels of the spacecraft, which includes a solar battery with four photovoltaic batteries installed on it, a device for rotating the solar panels, an amplifying-converting device, a control unit for the orientation of the solar panels towards the Sun, a block turning solar panels to a given position, two current regulators, a battery pack, a battery charger, a command generation unit for charging batteries, a load current sensor, a power supply system control unit, a power supply bus, a unit for measuring the density of the current flux of solar electromagnetic radiation, a block for determining solar activity, a block for determining the moment of time of the impact of particles on a spacecraft, a block for measuring the flux density of high-energy particles, a block for determining the moment of time of the beginning of control of solar batteries by load currents, a block of control of solar batteries by load currents, while the solar battery through its first the output that combines the outputs of two photovoltaic batteries is connected to the first input of the solar panel rotation device, and through the second output that combines the outputs of two other photovoltaic batteries, it is connected to the second input of the solar panel rotation device, and the outputs of the solar panel orientation control units towards the Sun and turning the solar panels to a given position are connected, respectively, to the first and second inputs of the amplification-converting device, the output of which, in turn, is connected to the third input of the solar panel rotation device, the first and second outputs of the solar panel rotation device are connected, respectively, to the inputs of the first and second regulators current, and the outputs of the current regulators are connected to the power supply bus of the spacecraft, the battery unit, with its input, through the battery charger, is connected to the power supply bus, while the battery charger is connected with its first input to the specified bus, and to the second input of the charger device for batteries, a load current sensor is connected, which is connected, in turn, to the power supply bus, the battery unit is connected with its output to the first input of the unit for generating commands for charging batteries, and the first output of the power supply system control unit is connected to the second input of the specified unit , the output of the unit for generating commands for charging batteries is connected to the third input of the battery charger, the second and third outputs of the power supply system control unit are connected to the first inputs of the control units for the orientation of solar panels towards the Sun and the rotation of solar panels to a given position, the third output of the device rotation of solar panels is connected to the second inputs of control units for the orientation of solar panels towards the Sun and rotation of solar panels to a given position, the output of the block for measuring the density of the current flux of solar electromagnetic radiation is connected to the input of the block for determining solar activity, the first output of which, in turn, is connected with the input of the block for determining the moment of time of the impact of particles on the spacecraft, the outputs of the block for determining the moment of time of the impact of particles on the spacecraft and the block for measuring the flux density of high-energy particles are connected to the first and second inputs of the block, respectively, for determining the moment of time of the start of control of solar panels by load currents, and the input of the block for measuring the flux density of high-energy particles is connected to the second output of the block for determining solar activity, the output of the block for determining the moment of time when solar panels start to be controlled by load currents is connected to the input of the power supply system control block, the fourth output of which, in turn, is connected to the first input of the block control of solar panels by load currents, the third input and output of which are connected to the third output of the solar panel rotation device and the third input of the amplifying-converting device, respectively, a block for determining the required current from solar panels, a block for determining the moments of time precursors of the negative impact of high-energy particles on a spacecraft and a unit for setting permissible values ​​of the battery charge level, while the first and second inputs and output of the unit for determining the required current from solar panels are connected to, respectively, the second output of the load current sensor, the second output of the battery charger and the second input of the solar battery control unit by load currents, the outputs of the unit for measuring the flux density of high-energy particles and the unit for measuring the density of the current flux of solar electromagnetic radiation are also connected to the corresponding

The essence of the proposed method is as follows.

The direct protective turn of the Security Council from the direction of the negative impact of high-energy particle flows is carried out when the density of high-energy particle flows exceeds certain specified threshold values. At the same time, as initial steps carried out before the direct implementation of protective measures, continuous monitoring of the current state of near-Earth space and current solar activity is carried out and the fulfillment and non-fulfillment of the criteria for a dangerous radiation situation, in particular the criteria for monitoring solar activity developed by the National Oceanic and Atmospheric Administration (NOAA), is analyzed ) (cm. ). In this case, situations where the criteria for unconditional danger have not yet been met, but the threshold of the previous level of danger has already been reached, should be considered as “precursor” situations of the negative impact under consideration.

When precursors of the negative impact of high-energy particle flows on the spacecraft appear, the maximum charge of the spacecraft SES AB is carried out. This makes it possible in the future, when the measured values ​​of the flux density of high-energy particles exceed the threshold values ​​compared with them, to turn the working surfaces of the SB panels away from the direction of the fluxes of these particles to the maximum possible angle, provided that the resulting shortage of electricity on board the spacecraft is compensated by the battery discharge. In this case, this value α s_min_AB of the angle of the protective flap SB is determined by the relation:

where I m is the maximum current generated when the illuminated working surface of the solar panels is oriented perpendicular to the sun's rays,

I SB - required current from SB.

In this case, the required current from the SB I SB is defined as the minimum required current that must be generated by the SB to provide consumers of the spacecraft, taking into account the possibilities of using the energy of the BAB SES of the spacecraft (i.e., when compensating for the emerging shortage of electricity on board the spacecraft due to the discharge of AB SES), based on the ratios:

where I n is the load current from spacecraft consumers,

I battery - the current maximum permissible discharge current of the battery of the SES spacecraft.

To implement the method, a system is proposed, shown in the drawing and containing the following blocks:

1 - SB, on the rigid substrate of the body of which four photovoltaic batteries are located;

2, 3, 4, 5 - BF 1, BF 2, BF 3, BF 4;

8 - BUOSBS;

9 - BRSBZP;

10, 11 - RT 1 and RT 2;

13 - ZRU AB;

14 - BFKZ AB;

16 - BUSES;

18 - BIPEMI;

20 - BOMVHF;

21 - BIPPCHVE;

22 - BOMVUSBTNZ;

23 - BUSBTNZ;

24 - block for determining the moments of time of the harbingers of the negative impact of high-energy particles on the spacecraft (BOMVPNVCH),

25 - block for determining the required current from solar panels (BOPTSB),

26 - block for setting permissible values ​​of the battery charge level (BZDZUZSB).

In this case, the SB (1) is connected through its first output, combining the outputs of BF 1 (2) and BF 4 (5), to the first input of the UPSB (6), and through the second output, combining the outputs of BF 2 (3) and BF 3 ( 5), connected to the second input of the UPSB (6). The outputs of BUOSBS (8) and BRSBZP (9) are connected, respectively, to the first and second inputs of the UPU (7), the output of which, in turn, is connected to the third input of the UPSB (6). The first and second outputs of the UPSB (6) are connected, respectively, to the inputs PT 1 (10) and PT 2 (11), and the outputs PT 1 (10) and PT 2 (11) are connected to the SE (17). The BAB (12) is connected to the SE (17) by its input through the AB (13) closed switchgear. In this case, the AB switchgear (13) is connected with its first input to the specified bus, and the accident output (15) is connected to the second input of the AB switchgear (13), the input of which is connected, in turn, to the ShE (17). The BAB (12) with its output is connected to the first input of the BFKZ AB (14), and the first output of the BUSES (16) is connected to the second input of the specified block. The output of the BFKZ AB (14) is connected to the third input of the ZRU AB (13). The second and third outputs of the BUSES (16) are connected, respectively, to the first inputs of the BUSBS (8) and BRSBZP (9). The third output of UPSB (6) is connected to the second inputs of BUOSBS (8) and BRSBZP (9). The BIPEMI output (18) is connected to the BOSA input (19). The first output of the BOSA (19) is connected to the input of the BOMVVCH (20). The outputs of BOMVVCH (20) and BIPPChVE (21) are connected to the first and second inputs of the BOMVUSBTNZ block (22), respectively. The input of the BIPPCHVE (21) is connected to the second output of the BOSA (19). The output of BOMVUSBTNZ (22) is connected to the first input of BUSES (16). BUSES (16) with its fourth output is connected to the first input of BUSBTNZ (23). The third output of the UPSB (6) is connected to the third input of the BUSBTNZ (23). The output of BUSBTNZ (23) is connected to the third input of the UPU (7). The first input of the BOPTSB (25) is connected to the second output of the DVT (15). The second input of the BOPTSB (25) is connected to the second output of the AB (13). The output of BOPTSB (25) is connected to the second input of BUSBTNZ (23). The output of the BIPPCHVE (21) is connected to the first input of the BOMVPNVCH (24). The output of the BIPEMI (18) is connected to the second input of the BOMVPNVCH (24). The output of the BOMVPNVCH (24) is connected to the second input of the BUSES (16). The first and second outputs of the BZDZUZSB (26) are connected to the third input of the BFKZ AB (14) and the fourth input of the ZRU AB (13), respectively.

The drawing also shows with a dotted line the mechanical connection of the UPSB (6) with the SB housing (1) through the output shaft of the battery drive.

In the spacecraft power supply mode, the system operates as follows. UPSB (6) serves for transit transmission of electricity from SB (1) to PT 1 (10) and RT 2 (11). Voltage stabilization on the SES power supply bus is carried out by one of the RTs. At the same time, the other RT is in a state with power transistors closed. Generators SB (1) (BF 1 - BF 4) operate in this case in short circuit mode. When the load power becomes greater than the connection power of the solar power generators (1), another RT switches to the voltage stabilization mode, and the energy of the unused generators is supplied to the power supply bus of the solar power plant. In certain periods, when the load power may exceed the power of the SB (1), the AB (13), due to the discharge of the AB unit (12), compensates for the shortage of electricity on board the spacecraft. For these purposes, the battery discharge regulator (13) serves as a battery discharge regulator, which, in particular, monitors the charge level of the battery and, upon reaching the minimum permissible value of the battery charge level, the value of which is supplied to the battery switchgear (13) from the BZDZUZSB (26), turns off BAB (12) from external load. In this case, the battery control switch (13), based on the current charge level of the battery, determines and supplies to its second output the current value of the permissible battery discharge current (in the mode of disconnecting the battery (12) from the external load, this value is zero).

In addition to the specified regulator, the battery charger (13) also contains a battery charge regulator. To carry out charge-discharge cycles in the AB (13), information from the DTN (15) is used. The charge of the BAB (12) is carried out by the ZRU AB (13) through the BFKZ AB (14). For the case of metal-hydrogen batteries, it is described in. The bottom line is that the density of hydrogen in the battery casing is determined using pressure sensors installed inside the batteries and temperatures on the battery cases. In turn, the density of hydrogen determines the charge level of the battery. When the hydrogen density in the battery drops below a set level, a command is issued to charge it, and when the maximum density level is reached, a command is issued to stop charging. The indicated battery charge levels are regulated by commands from the BFKZ AB (14), while the values ​​of the maximum permissible charge level of the battery are supplied to the BFKZ AB (14) with the BZDZUZSB (26). Maintaining the batteries in a maximally charged state negatively affects their condition, and the batteries are maintained in the current self-discharge mode, in which the operation of charging the batteries is performed only periodically (for example, when controlling the SES of the Yamal-100 spacecraft - once every few days, when the charge level decreases BAB at 30% of the maximum level).

Simultaneously with operation in the spacecraft power supply mode, the system solves the problem of controlling the position of the planes of the solar panel panels (1).

Upon command from the BUSES (16), the BUSBS block (8) controls the orientation of the SB (1) to the Sun. BUOSBS (8) can be implemented on the basis of a spacecraft VESSEL (see). In this case, the input information for the satellite control algorithm is: the position of the unit direction vector to the Sun relative to the coordinate axes associated with the spacecraft, determined by the algorithms of the kinematic contour of the vessel; the position of the SB relative to the spacecraft body, obtained in the form of the current measured values ​​of the angle α with the UPSB remote control (6). The output information of the control algorithm is commands to rotate the SB relative to the axis of the output shaft of the UPSB (6), commands to stop rotation. The UPSB remote control (6) produces discrete signals about the position of the SB (1).

BIPEMI (18) measures the current fluxes of solar EMR and transmits them to BOSA (19). In BOSA (19), by comparing current values ​​with given threshold values, the onset of solar activity is determined. According to the command coming from the first output of the BOSA (19) to the input of the BOMVVCH (20), in the indicated last block the moment in time of the possible beginning of the impact of high-energy particles on the spacecraft is determined. From the second output of the BOSA (19) through the input of the BIPPCHVE (21), a command is issued to begin measuring the flux density of high-energy particles.

From the output of BIPPChVE (21), the measured value of the flux density of high-energy particles is transmitted to the first input of BOMVPNVP (24) and to the second input of BOMVUSBTNZ (22). The measured values ​​of the current solar EMR fluxes are supplied to the second input of the BOMVPNVCH (24) from the output of the BIPEMI (18).

BOMVPNVCh (24) assesses the dynamics of changes in the flux density of high-energy particles and identifies situations that can be considered as harbingers of the negative impact of particles on the spacecraft. Such situations are when the measured flux density of high-energy particles exceeds specified critical values ​​and there is a tendency for its further increase. When identifying and identifying such situations, solar EMR flux data obtained from BIPEMI is also used (18). When registering such precursor situations in the BOMVPNVCh (24), a signal is generated at the output of this block and sent to the second input of the BUSES (16).

Upon command at the second input of the BUSES (16), this unit sends a command to the BFKZ AB (14), according to which this unit, through the closed switchgear AB (13), charges the BAB (12) to the maximum charge level. At the same time, for the case of metal-hydrogen batteries (see), using pressure sensors installed inside the batteries and temperatures on the battery cases, the density of hydrogen in the battery case is determined, from which the charge level of the battery is determined. When the maximum density level is reached, a command is issued to stop charging.

The inputs of the BOPTSB (25) from the second outputs of the DTN (15) and the closed switchgear battery (13) receive the current values ​​of the load current from the consumers of the spacecraft I n and the permissible discharge current of the battery I AB. Using these values ​​of BOPTSB (25), using relations (4), (5) determines the value of I SB - the current minimum permissible value of the required current from the SB (taking into account the possibility of consumers using energy from the BAB (12)), and outputs it to the second input BUSBTNZ (23).

Information about the time of the possible start of the impact of particles on the spacecraft is transmitted from the output of the BOMVVCH (20) to the BOMVUSBTNZ (22) through its first input. In BOMVUSBTNZ (22), the actual assessment of the negative impact of FVS is carried out by comparing the current measured value of the impact characteristic with threshold values, starting from the time point determined by BOMVUSBTNZ (20). A necessary condition for receiving a command at the output of BOMVUSBTNZ (22) is the presence of two signals - from the outputs of BOMVVCH (20) and BIPPCHVE (21).

When BOMVUSBTNZ (22) issues a command to the first input of BUSES (16), this block generates a command at its fourth output, which connects to the control of SB BUSBTNZ (23).

BUSBTNZ (23) determines the angle α s_min_AB by expression (3). To calculate the specified angle, the current value of the required current from the SB, obtained from the BOPTSB (25), is used. In addition, from the UPSB remote control (6) the specified block receives information about the current value of the SB rotation angle α. Having determined the value of the angle α s_min_AB, the algorithm embedded in BUSBTNZ (23) compares it with the current value of the angle α and calculates the mismatch angle between α and α s_min_AB and the required number of control pulses to activate the control drive SB (1). Control pulses are transmitted to the control unit (7). After converting and amplifying the indicated pulses in the UPU (7), they arrive at the input of the UPS (6) and set the drive in motion.

When BOMVUSBTNZ (22) does not issue a command to the first input of BUSES (16), this block, depending on the spacecraft flight program being executed, transfers control of the SB (1) to one of the blocks BUOSBS (8) and BRSBZP (9).

The functioning of the BUSBS (8) is described above.

BRSBZP (9) controls SB (1) according to program settings. The SB control algorithm (1) according to software settings allows you to install the battery in any specified position α=α z . In this case, to control the rotation angle in the BRSBZP (9), information from the UPSB remote control (6) is used.

The implementation of BOMVUSBTNZ (22) and BOMVPNVCh (24) is possible both on the basis of hardware and software of the spacecraft control center and on board the spacecraft. At the outputs of BOMVUSBTNZ (22) and BOMVPNVCH (24), the commands “start control of the power supply based on load currents” and “start control of the solar power system in preparation mode for the negative impact of high-energy particles on the spacecraft” are formed, respectively, which are sent to BUSES (16), when In this case, the last command is functionally perceived by BUSES (16) as a command to charge the battery to the maximum charge level.

An example of the implementation of BUSES (16) can be the radio means of the service control channel (SCU) onboard systems of the Yamal-100 spacecraft, consisting of an earth station (ES) and on-board equipment (BA) (see description in). In particular, the BA SKU together with the GS SKU solves the problem of issuing digital information (DI) to the on-board digital computer system (OBDS) of the spacecraft and its subsequent acknowledgment. BTsVS, in turn, controls the blocks BUOSBS (8), BRSBZP (9), BUSBTNZ (23), BFKZ AB (14).

In this implementation of BUSES (16), the interaction of the SKU BA in terms of data exchange is carried out via the main exchange channel (MEC) in accordance with the MIL-STD-1553 interface. As a subscriber of the BCWS, a device is used - an interface unit (UB) from the BA SKU. The BCWS processor periodically polls the BS state to determine the availability of a data packet. If the packet is available, the processor begins data exchange.

UPU (7) plays the role of an interface between BUOSBS (8), BRSBZP (9), BUSBTNZ (23) and UPSB (6) and serves to convert digital signals into analogue ones and amplify the latter.

BUSBTNZ (23) is the on-board unit of the spacecraft, commands to which come from BUSES (16). The implementation of BUSBTNZ (23), BOPTSB (25), BZDZUZSB (26) can be carried out on the basis of the spacecraft BTsVS (see,).

Thus, an example of the implementation of the fundamental blocks of the system is considered.

Let us describe the technical effect of the proposed inventions.

The proposed technical solutions provide a reduction in the negative impact of high-energy particle flows on the working surface of the solar system at the moments when the “protective” lapel of the solar panel is performed from the direction towards the Sun. This is achieved by reducing the area of ​​the working surface of the SB, which is negatively affected by the flows of these particles, by maximizing the angle of the normal to the working surface of the SB from the direction towards the Sun, while ensuring that the requirement of providing the spacecraft with electricity is met. Maximization of the turning angle is achieved by the fact that the solar power system of the spacecraft is previously brought into a state of maximum charge of the battery, which makes it possible to implement the maximum possible angle of the “protective” turning of the solar cell from the direction towards the Sun. Considering, for example, that when controlling the SES of the Yamal-100 spacecraft after the operation of charging the battery to the maximum level, the increase in the possible discharge current of the battery is about 30%, then a corresponding increase in the angle of the “protective” flap of the battery and, as a consequence, a decrease in the negative impact of particle flows high energies on the working surface of the SB is a significant value.

LITERATURE

1. Eliseev A.S. Space flight technology. Moscow, "Mechanical Engineering", 1983.

2. Rauschenbach G. Handbook for the design of solar panels. Moscow, Energoatomizdat, 1983.

3. Flight rules during joint operations of the SHUTTLE and the ISS. Tom S. Flight Operations Directorate. Space Center named after Lyndon B. Johnson. Houston, Texas, main version, 11/8/2001.

4. Spacecraft power supply system. Technical description. 300GK.20Yu. 0000-ATO. RSC Energia, 1998.

5. Center B.I., Lyzlov N.Yu., Metal-hydrogen electrochemical systems. Leningrad. "Chemistry", Leningrad branch, 1989.

6. Spacecraft motion control and navigation system. Technical description. 300GK.12Yu. 0000-ATO. RSC Energia, 1998.

7. Galperin Yu.I., Dmitriev A.V., Zeleny L.M., Panasyuk L.M. The influence of space weather on the safety of aviation and space flights. "Flight 2001", pp. 27-87.

8. Engineering reference book on space technology. Publishing house of the Ministry of Defense of the SSR, M., 1969.

9. Grilikhes V.A., Orlov P.P., Popov L.B. Solar energy and space flights. Moscow, "Science", 1984.

10. Earth station of the service control channel of the Yamal spacecraft. Manual. ZSKUGK.0000-ORE. RSC Energia, 2001.

11. Onboard equipment of the service control channel of the Yamal spacecraft. Technical description. 300GK.15Yu. 0000A201-OTO. RSC Energia, 2002.

12. Kovtun V.S., Solovyov S.V., Zaikin S.V., Gorodetsky A.A. A method for controlling the position of solar panels of a spacecraft and a system for its implementation. RF patent 2242408 according to application 2003108114/11 dated March 24, 2003

1. A method for controlling the position of the solar panels of a spacecraft, including turning the solar panels into a working position that ensures the supply of electricity to the spacecraft and corresponding to the alignment of the normal to their illuminated working surface with the plane formed by the axis of rotation of the solar panels and the direction to the Sun, measuring the density of the current flux of solar electromagnetic radiation, determining the moment in time when solar activity begins, determining the moment in time when high-energy particles reach the surface of the spacecraft, measuring the flux density of high-energy particles, comparing the measured values ​​of the flux density of high-energy particles with threshold values, turning solar panels at an angle between the normal to their illuminated working surface and direction towards the Sun, corresponding to the minimum area of ​​influence of high-energy particle fluxes on the surface of solar panels while simultaneously providing the spacecraft with electricity, at the moment of time when the measured values ​​of high-energy particle flux density exceed threshold values ​​and the solar panels return to their operating position at the point in time at which the density of high-energy particle fluxes becomes below threshold values, characterized in that they additionally determine the moments in time when the precursors of the negative impact of high-energy particle fluxes on the spacecraft appear and at the specified times the batteries of the spacecraft power supply system are charged to the maximum charge level, if the measured values ​​of the flux density of high-energy particles exceed the threshold values ​​compared with them, the solar panels are rotated until the angle between the normal to their illuminated working surface and the direction to the Sun α s_min_AB is reached, corresponding to the minimum area of ​​influence of the fluxes of high-energy particles on the surface of solar panels, while simultaneously providing the spacecraft with electricity from solar and rechargeable batteries of the power supply system, and determined by the ratio

α s_min_AB =arccos (max(0, I n -I AB )/I m),

where I n is the load current of the spacecraft consumers;

I m - maximum current generated when the illuminated working surface of solar panels is oriented perpendicular to the sun's rays;

I AB - the current permissible discharge current of the rechargeable batteries, and the resulting shortage of electricity on board the spacecraft is compensated by discharging the rechargeable batteries, while monitoring the charge level of the rechargeable batteries and, upon reaching the minimum permissible value of this level, the current value of the permissible discharge current of the rechargeable batteries is reset and disconnecting batteries from external load.

2. A system for controlling the position of the solar panels of the spacecraft, which are four photovoltaic solar panels mounted on panels, including a device for rotating the said solar panels, an amplifying-converting device, a control unit for the orientation of the solar panels towards the Sun, a unit for turning the solar panels to a given position, two current regulators, a battery pack, a battery charger, a command generation unit for charging batteries, a load current sensor, a power supply system control unit, a power supply bus, a unit for measuring the density of the current flux of solar electromagnetic radiation, a solar activity detection unit, a determination unit moment of time of impact of high-energy particles on the spacecraft, a unit for measuring the flux density of high-energy particles, a unit for determining the moment of time of the beginning of control of solar batteries by load currents, a unit of control of solar batteries by load currents, while the solar battery through its first output, combining the outputs of two photovoltaic batteries, is connected to the first input of the solar panel rotation device, and through the second output, which combines the outputs of two other photovoltaic batteries, is connected to the second input of the solar panel rotation device, and the outputs of the control units for the orientation of solar panels towards the Sun and the rotation of solar panels to a given position are connected, respectively, to the first and second inputs of the amplification-converting device, the output of which, in turn, is connected to the third input of the solar panel rotation device, the first and second outputs of the solar panel rotation device are connected, respectively, to the inputs of the first and second current regulators, and the outputs of the current regulators are connected to the power supply bus of the spacecraft, the battery unit is connected with its input, through the battery charger, to the power supply bus, while the battery charger is connected with its first input to the specified bus, and to the second input of the battery charger batteries, a load current sensor is connected, which is connected, in turn, to the power supply bus, the battery block is connected with its output to the first input of the block for generating commands for charging batteries, and the first output of the power supply system control unit is connected to the second input of the specified block, the output of the block generating commands to charge the batteries is connected to the third input of the battery charger, the second and third outputs of the power supply system control unit are connected to the first inputs of the control units for the orientation of the solar panels towards the Sun and the rotation of the solar panels to a given position, the third output of the solar panels rotation device connected to the second inputs of the control units for the orientation of solar panels towards the Sun and the rotation of solar panels to a given position, the output of the block for measuring the density of the current flux of solar electromagnetic radiation is connected to the input of the block for determining solar activity, the first output of which, in turn, is connected to the input of the block determining the moment of time of the impact of particles on the spacecraft, the outputs of the block for determining the moment of time of the impact of particles on the spacecraft and the block for measuring the flux density of high-energy particles are connected to, respectively, the first and second inputs of the block for determining the moment of time of the start of controlling solar panels by load currents, and the input the block for measuring the flux density of high-energy particles is connected to the second output of the block for determining solar activity, the output of the block for determining the moment of time when solar panels start to be controlled by load currents is connected to the input of the power supply system control block, the fourth output of which, in turn, is connected to the first input of the control block solar panels according to load currents, the third input and output of which are connected to, respectively, the third output of the solar panel rotation device and the third input of the amplifying-converting device, characterized in that it additionally includes a block for determining the required current from solar panels, a block for determining the moments of occurrence harbingers of the negative impact of high-energy particles on the spacecraft and the unit for setting permissible values ​​of the battery charge level, while the first and second inputs and output of the unit for determining the required current from solar panels are connected to, respectively, the second output of the load current sensor, the second output of the battery charger batteries and the second input of the solar panel control unit for load currents, the outputs of the unit for measuring the flux density of high-energy particles and the unit for measuring the density of the current flux of solar electromagnetic radiation are connected

The invention relates to astronautics and can be used in space activities - research of outer space, planets of the solar system, observations of the Earth from space, etc., in which it is necessary to determine the spatial coordinates of spacecraft (SV) and the components of its velocity vector.

The invention relates to rocket and space technology and can be used in the creation of launch vehicles (LVs), including conversion ones, for launching spacecraft into low-Earth orbits.

The invention relates to the field of space technology, namely to power supply systems for spacecraft, and can be used to control the position of their solar panels