RU2797430C1 - Device for resetting the kinetic moment and controlling the altitude of the spacecraft using a magnetic system - Google Patents

Abstract

FIELD: space engineering.

SUBSTANCE: invention relates to devices for controlling the orientation of a spacecraft by its interaction with the geomagnetic field, namely for removing the kinetic moment of inertial executive bodies. The device includes a block of electromagnets with inductors, a micro-computer, a block of reversible switches and a control device for reversible switches, a block of current sensors, a block of DC voltage converters, a block of angular velocity sensors, a block of flywheel motors and a magnetometer. In the magnetic system, electromagnets ensure the formation of high values of magnetic moments along the associated X, Y, Z axes of the spacecraft.

EFFECT: increased efficiency and reliability of the magnetic system.

1 cl, 4 dwg

Description

Purpose

The present invention relates to the magnetic system of a spacecraft (SC) formed by interaction with the Earth's geomagnetic field. inertial executive bodies in the form of flywheel engines (DM) in the orientation system.

State of the art

The most widely known systems that create an external mechanical torque on the spacecraft are systems that use executive bodies with gas-jet nozzles or microjet engines (see, for example, patents, RF, No. 2271317, 2648906). These executive bodies, throwing cold or hot gas through a nozzle into outer space, create a force acting on the apparatus.

However, they have a significant drawback - the consumption of the working fluid, as a result of which the useful time of the operation of the spacecraft is limited (in modern spacecraft for remote sensing of the Earth, increased requirements are imposed on the reliability of operation, taking into account the service life of the spacecraft, as a rule, at least 7-10 years) and, in addition, for spacecraft, these systems are practically unacceptable due to the presence of a working fluid that increases the mass and dimensions of the spacecraft.

Therefore, for spacecraft, the most suitable control devices are, for example, flywheel motors (DM), often used in spacecraft orientation and stabilization systems (see, for example, patent, RF, No. 1839928), which are distinguished by high stabilization accuracy, small size, ease of control , as well as the absence of consumption of non-renewable working fluid. Flywheel motors create an internal control torque (create a mechanical moment around an axis parallel to the axis of rotation of the flywheel motor rotor), but they cannot fend off external disturbances.

They are inertial carriers of the kinetic moment, so one of the disadvantages of engines of this type is the need to reset the kinetic moment of the flywheels when they reach the maximum speed of rotation. In the event that some external moment acts on the body of the spacecraft (aerodynamic drag forces, light pressure, gravitational field or other external forces), then the body acquires some angular velocity around some axis over a certain time. This "parasitic" speed can be eliminated only by an external moment (see A.G. Iosifyan, Electromechanics in space. "Cosmonautics, astronomy" No. 3. 1977). Physically, this means that the "parasitic" rotation is suspended if the kinetic moment received by the body is transferred "inside", starting, for example, the rotor of the flywheel motor in the direction along which the external force and external torque acted, while resetting the kinetic moment of the spacecraft .

Known systems that create an external mechanical moment on the spacecraft, which do not require the presence of any expendable working fluid - these are magnetic systems operating in near space under the influence of the Earth's geomagnetic field (altitude up to approximately 9000 km).

As executive bodies in these magnetic systems, electromagnets are used that form magnetic moments, which, interacting with the geomagnetic field of the Earth, create external control mechanical moments on the spacecraft (see A.P. Kovalenko. Magnetic control systems for spacecraft. M., "Engineering ”, 1975, p. 9), capable of providing the process of calming down during the separation of the spacecraft from the launch vehicle or upper stage, as well as, with the combined use of a magnetic system with flywheel engines, the release of kinetic momentum and attitude control (movement around the center of mass) KA.

Thus, when creating modern spacecraft, it is very important to develop a reliable high-performance magnetic system that does not require a working fluid, with optimal energy costs, which makes it possible to ensure the reset of the kinetic moments of inertial actuators, the orientation and stabilization of spacecraft with a given accuracy, as well as the process of calming during separation of the spacecraft from a launch vehicle or upper stage.

In the general case, the control moment M is determined by the basic control equation (see A.P. Kovalenko. Magnetic control systems for spacecraft. M., Mashinostroyeniye. 1975, p. 21):

where Q is the vector of the generated magnetic moment by the magnetic system;

B is the vector of magnetic induction of the Earth's magnetic field.

It should be noted that in the magnetic system, a control moment cannot be created relative to the direction coinciding with the magnetic induction vector of the Earth's magnetic field, and independent moments simultaneously relative to all control axes.

Known is the "Method of magnetic unloading of spacecraft flywheel engines" (patent RF, No. 2568827) in which, when the critical level of the accumulated kinetic moment is exceeded by the flywheel engines, the spacecraft is deployed using a DM around two mutually perpendicular axes of the spacecraft in such a way that the DM axis with a smaller accumulated angular momentum coincided with the magnetic induction vector of the external magnetic (geomagnetic) field. The magnetic system is arrested and the moment arising from the interaction of the magnetic system and the external magnetic field is used to reset the kinetic moment of the DM lying in a plane perpendicular to the magnetic induction vector of the external magnetic field. After the angular momentum has been reset, the magnetic system is unlocked and the spacecraft, if necessary, is returned to the reference coordinate system, or a new turn takes place in order to reset the DM angular momentum, the axis of which coincided with the magnetic induction vector of the external magnetic (geomagnetic) field.

The disadvantage of this invention is that mechanical stoppers (cages) are required for caging and re-caging the magnetic system relative to the spacecraft body, which leads to a significant structural complication, especially for large spacecraft, and, in general, to a decrease in the reliability of the magnetic system.

In addition, the turn of the spacecraft is provided with the help of DM requiring unloading, which are close to saturation. This is an unfavorable mode, which reduces the reliability of the magnetic system, because The DM can act on the spacecraft only when the speed of rotation of the DM rotor changes. When the rotation speed reaches the limit, the DM loses the ability to influence the spacecraft.

Known "Method for resetting the kinetic moment of the inertial actuators of the spacecraft and a device for implementing the method" (patent, RF, No. 2625687), the device of which includes sensors of the kinetic moment components along the three coordinate axes X, Y, Z, sensors of the geomagnetic induction components along three coordinate axes X, Y, Z; the inputs of which are connected to three outputs of the control unit, three comparators, the first inputs of which are connected to the outputs of the DAC. and the threshold signal - σ _o is fed to the second inputs, three elements And, the first inputs of which are connected to the outputs of the comparators, and the second inputs are connected to the fourth output of the control unit, the three electronic keys, the inputs of which are connected to the outputs of the DAC, the control inputs of the electronic keys are connected to the outputs of the elements And, and the outputs are connected to the inputs of the magnetic actuators containing series-connected reversing switches and electromagnets.

The disadvantage of this invention is the low reliability of the device, in view of the fact that the period of active existence of modern spacecraft is at least 7 years and in case of failure, for example, electromagnets and a number of other key electronic devices, there is no possibility to maintain the operability of the magnetic system.

In addition, the technical support in each coil X, Y, Z of the magnetic executive body of the current strength in accordance with the values of the kinetic moment K _x , K _y , K _z , (σ _x , σ _y , σ _z ) is carried out by electronic microcircuits of small integration, which , due to the low functionality under conditions of limitation in terms of mass and dimensions, have a low accuracy in the formation of the required current strength, corresponding to the value of the kinetic moment.

Closest to the proposed invention is "Device for the formation of a mechanical moment by an electromagnet in the magnetic orientation system of a spacecraft" (patent, RF, No. 2672316), taken by the authors as a prototype.

This device for generating a mechanical moment by an electromagnet in the spacecraft's magnetic orientation system includes an electromagnet with an inductance coil and an electronic device containing a control device and a bridge circuit on semiconductor power elements connected to the outputs of the control device, in addition, between the electromagnet and the electronic device a relay unit with a connector for connecting to an electromagnet, wherein the electromagnet winding consists of two inductors - the main and the backup, and the electronic device contains two identical channels - the main and the backup, each of which, in addition to the bridge circuit and the control device, contains a micro-computer and two optocoupler took into account, parallel back-to-back LEDs of which with one output are connected to one point of the diagonal of the bridge circuit, the other - to the relay node, which is also connected to the second point of the diagonal of the bridge circuit, while the group of inputs of the control device is connected to the group of outputs of the micro-computer. a separate output of which is connected to the input of the relay node, and separate inputs - to the outputs of the photodetector of the optocoupler nodes.

The disadvantage of the prototype is the low efficiency of the control of the magnetic system, especially for large spacecraft (mass more than 1000 kg), and there are no conditions for improving the reliability of the key devices of the device in order to achieve the required active lifetime for modern spacecraft. In accordance with GOST R 56526-2015, reliability is a set of properties that characterize the ability of a spacecraft to ensure, in the process of operation, obtaining the output effect specified in the terms of reference under given conditions and operating modes. The main properties of reliability are non-failure operation, durability, and persistence.

The high control efficiency of the magnetic system should be understood as the formation of a high control magnetic moment, while conditions should be provided for optimally low power consumption by the actuating output bodies - electromagnets, which mainly determine the maximum power consumption of the magnetic system from the onboard power source of the spacecraft.

The need to form a high control magnetic moment is due to the fact that the maximum value of the mechanical moment acting on the spacecraft is formed under the condition that the vector of the magnetic moment of the electromagnet and the vector of the induction of the Earth's geomagnetic field are perpendicular and vice versa, the minimum value of the mechanical moment, reaching a complete absence - when they are parallel . This is due to the fact that the main magnetic field of the Earth is determined by sources located in the liquid core and at the core-mantle boundary. The world model of the normal field, based on the representation of the normal field as a series of spherical functions, the coefficients of which are determined every 5 years based on the global network of magnetic observatories and is described by the expressions:

where X, Y, Z are the northern, eastern and vertical components, respectively;

R is the average radius of the Earth, r is the distance from the point to the center of the Earth;

λ - longitude, θ=π/2-ϕ - addition to latitude;

- присоединенная функция Лежандра первого рода;

- associated Legendre function of the first kind;

- сферические гармонические коэффициенты.

- spherical harmonic coefficients.

The geomagnetic field of the Earth is characterized by the presence of both global and regional anomalies and is continuously changing in time, i.e. characterized by the so-called secular course (see RF patent, No. 2447405). In connection with this, the angle between the vector of the magnetic moment of the electromagnet and the vector of the induction of the geomagnetic field of the Earth constantly changes when the spacecraft moves in orbit, therefore, in order to increase the efficiency of the magnetic system, it is necessary to reduce the turn-on time of the actuators, first of all, by increasing the control torque on the spacecraft.

Reducing the power consumption of electromagnets is due to the most important requirement for all consumers of electricity from the onboard power supply of the spacecraft, which includes a solar battery that converts the constantly changing light energy of the Sun, and a storage battery, which is one of the most critical links in the power supply system of the spacecraft, as well as control and control (see, for example, patents No. 2689401, No. 2724111).

The aim of the invention is to improve the efficiency and reliability of the magnetic system.

Disclosure of invention

The essence of the proposed device for resetting the kinetic moment and controlling the attitude of the spacecraft using a magnetic system lies in the technical support for increasing the reliability of the magnetic system and the efficiency of controlling the spacecraft (increasing the control torque), as well as reducing the energy consumption of the actuating output elements - electromagnets, which in the magnetic system are mainly determine the maximum power consumption of electricity from the onboard power source of the spacecraft.

The increase in the reliability of the operation of the magnetic system and the efficiency of control of the spacecraft is ensured by the fact. that the electromagnets along the axes of the spacecraft EM _X , EM _Y , EM _Z in the present invention are made of n electromagnets (n≥1), each of which contains two windings L1 and L2, operating in hot mode. These n electromagnets, which form the magnetic moments of one direction, can be connected in parallel or in series and are located along the associated X, Y, Z axes of the spacecraft.

Reducing the power consumption of electromagnets is carried out by forming microcontrollers with high accuracy of voltages (respectively, current strength) in electromagnets EM _X , EM _Y , EM _Z in accordance with the values of the kinetic moment K _x , K _y , K _z along the associated axes of the spacecraft.

In addition, the high reliability of the components of the device for resetting the angular momentum and controlling the attitude of the spacecraft using a magnetic system is ensured through the reasonable integrated and cumulative use of redundancy, redundancy, and majority voting.

The device for resetting the angular momentum and controlling the attitude of the spacecraft using a magnetic system includes an electromagnet with an inductance coil consisting of two inductors - a main and a backup one, a micro-computer, a reversing switch made according to a bridge circuit on semiconductor power elements, a relay assembly, two current sensors and a control device for reversing switches.

Introduction to the device for resetting the angular momentum and controlling the attitude of the spacecraft using the magnetic system of a block of electromagnets containing electromagnets installed along the associated X, Y, Z axes of the spacecraft, respectively, electromagnets EM _x , EM _y , EM _z , each of which consists of n ( n≥1) electromagnets, a block of reversible switches, consisting of three identical reversing switches along the associated axes X, Y, Z of the spacecraft, a block of current sensors, consisting of three identical optocoupler current sensors along the associated axes X, Y, Z of the spacecraft, a block converters of direct voltage, consisting of three identical converters of direct voltage along the connected X, Y, Z axes of the spacecraft, the block of sensors of angular speeds, the block of motor-flywheels and the magnetometer allows to increase the efficiency and reliability of the magnetic system.

By introducing into the device a block of electromagnets containing electromagnets EM _x , EM _y , EM z installed along the associated axes X, Y, _Z , respectively, each of which consists of n (n≥1) electromagnets, it is possible to increase the magnetic moments formed by the electromagnets , which allows the use of the magnetic system, including for large spacecraft with high efficiency of spacecraft control (control of high mechanical torque).

The introduction of a block of reversible switches and a block of current sensors into the device ensures the supply of the required voltages for electromagnets along the connected axes X, Y, Z of the spacecraft, as well as the direction of the current in them, which determines the direction of the mechanical control moment on the spacecraft.

The introduction of a block of DC voltage converters into the device ensures a reduction in the power consumption of electromagnets due to the formation of voltages (respectively, current strength) in the electromagnets along the associated axes of the spacecraft EM _X , EM _Y , EM _Z in accordance with the values of the kinetic moment K _x , K _y , K _z (σ _x , σ _y , σ _z ) with high accuracy. High accuracy is ensured by electronic components with a high degree of integration and high functionality - microcontrollers that "work" in DC voltage converters along the associated X, Y, Z axes of the spacecraft. Microcontrollers generate PWM signals that drive high-frequency inverters with rectifiers at the output.

Graphic illustrations

The invention is illustrated by the graphical figures of FIG. 1 and FIG. 2.

The given graphic figure (Fig. 1) shows a block diagram for the implementation of the proposed device for resetting the angular momentum and controlling the attitude of the spacecraft using a magnetic system containing components indicated by positions:

• BEM - a block of electromagnets (along the associated axes X, Y, Z of the spacecraft, respectively, electromagnets EM _x , EM _y , EM _z ) - 1;

• main inductors of the electromagnet - L1-1, L1-2…, L1-n;

• backup inductors of the electromagnet - L2-1, L2-2…, L2-n;

• GPZ (geomagnetic field of the Earth) - 2;

• mechanical moments formed along the associated axes X, Y, Z of the spacecraft (M _x , M _y , M _z ) - 3;

• spacecraft body - 4;

• block DM (block of sensors-flywheels) - 5;

• BDUS (unit of angular velocity sensors) - 6;

• magnetometer - 7;

• microcomputer - 8;

• DBK - a block of reversible switches (along the associated axes X, Y, Z of the spacecraft, respectively 9 _x , 9 _y , 9 _z ) - 9;

• block of current sensors (along the X, Y, Z axes of the spacecraft, respectively 10 _x , 10 _y , 10 _z )-10;

• optocoupler units of current sensors along the associated X-axis of the spacecraft - 10 _x -1, 10 _x -2;

• relay node -11;

• CU RK (control device for reversing switches) - 12;

• BC (on-board computer) - 13;

• NKU (ground control complex) - 14;

• PPV - block of DC voltage converters (on the connected X, Y, Z axes of the spacecraft, DC voltage converters, respectively, DC _x - 15 _x , DC _y - 15 _y , DC _z - 15 _z ) - 15;

• MK (microcontroller) - 16;

• And (inverter) - 17;

• rectifier - 18;

• SC UPS (SC onboard power supply) - 19.

Between the on-board computer BC 13 and the micro-computer 8 is the exchange of command and software information, navigation data.

For all modern spacecraft, the exchange is carried out via a multiplex exchange channel (requirement of GOST R 52070-2003).

Telemetric information signals from the on-board computer BC 13 are transmitted to the ground control center of the NKU 14 via telemetry (TM), and from the ground control center of the NKU 14 to the on-board computer BC 13 - control commands (KU).

In FIG. Figure 2 shows the graphs of the formation of useful and parasitic components of mechanical moments in the magnetic system.

Implementation of the invention

The device for resetting the angular momentum and controlling the attitude of the spacecraft using a magnetic system includes an electromagnet with an inductance coil consisting of two inductors - the main L1-1 and the backup L2-1, micro-computer 8, reversing switch 9 _x , made according to bridge circuit on semiconductor power elements, and a control device for reversing switches UU RK 12, connected by an output to the control inputs of a reversing switch 9 _x , one diagonal point of which is connected through a relay node 11 to one end of the main inductance coil of the electromagnet L1-1, the second end of which is also through the relay node 11 it is connected to one output, included in the block of current sensors 10, two parallel counter-connected, LEDs of two optocoupler nodes of current sensors 10 _x -1 and 10 _x -2, the second output of which is connected to the second point of the diagonal of the reversing switch 9 _x , at the same time, a separate output of the micro-computer 8 is connected to the input of the control device for reverse channels CU RK 12, and separate inputs are connected to the outputs of the photodetector (F) of optocoupler units of current sensors 10 _x -1 and 10 _x -2.

The device additionally includes a block of electromagnets BEM 1, a block of reversible switches BRK 9, a block of current sensors 10, a block of DC voltage converters 15, a block of angular velocity sensors (BDUS) 6, a block of flywheel motors 5 and a magnetometer 7, while the block of electromagnets BEM 1 contains installed along the associated axes X, Y, Z of the spacecraft, respectively, electromagnets EM _x , EM _y , EM _z , each of which consists of n (n≥1) electromagnets connected in series or in parallel through a relay node 11 with a block of reverse switches DBK 9 and containing main L1-1, L1-2…, L1-n and backup L2-1, L2-2…, L2-n inductors operating in hot mode;

block of reversing switches DBK 9, consisting of three identical reversing switches along the associated axes X, Y, Z of the spacecraft 9 _x , 9 _y , 9 _z , each of which is connected to the outputs of the control device for reversing switches UU RK 12 and the outputs of the block of DC/DC converters 15, at the same time, each reversing switch 9 _x , 9 _y , 9 _z is connected to identical devices of optocoupler nodes 10 _x , 10 _y , 10 _z of the current sensor unit 10, respectively, along the associated axes X, Y, Z of the spacecraft, which are connected with a relay node 11 and a group of individual inputs of the microcomputer 8;

block of angular velocity sensors BDUS 6, block DM 5 and magnetometer 7 are connected to a group of separate inputs of microcomputer 8. which is connected by a group of information outputs to the input of the control device for reversing switches UU RK 12, inputs of the DC voltage converter block 15 and microcontrollers 16 in each DC voltage converter PPN _x 15 _x , PPN _y 15 _y , PPN _z 15 _z , while. The PWM outputs of the microcontrollers 16 are connected to the control inputs of the inverters 17, the outputs of which, through the rectifiers 18, are the outputs of the DC/DC converter unit 15, and are connected via the power buses to the onboard power supply of the UPS KA 19;

components of magnetic moments Q _x , Q _y , Q _z formed by the block of electromagnets BEM 1 interacting with the components of magnetic inductions B _x , B _y , B _z of the geomagnetic field of the Earth GPZ (2) form the control moments M _x , M _y , M _z (3) , affecting the body of the spacecraft 4, while the components of the magnetic induction B _x , B _y , B _z of the geomagnetic field of the Earth GPZ (2) are measured by the magnetometer 7.

Description of the device for resetting the kinetic moment and controlling the attitude of the spacecraft using a magnetic system

The executive body of the magnetic system is an electromagnet, which consists of an inductor with a core placed inside. When direct current is passed through the turns of the inductor, the core is magnetized and the electromagnet acquires the properties of a permanent magnet that forms a magnetic field that disappears when the current stops flowing, while the core is demagnetized.

Electromagnets EM _X , EM _Y , 3M _Z of the block of electromagnets BEM 1 are placed on the body of the spacecraft for mechanical control around the center of mass, as a rule, along orthogonal coupled axes X, Y, Z of the spacecraft.

In the proposed invention, reliable electromagnets with a rod-type core made of a soft magnetic material with copper wire windings are used as electromagnets. For example: two inductors are wound with copper wire on a rod core of a round section made of a soft magnetic material (for example, to increase reliability, a hot backup mode is used in the operation of inductors L1-1 and L2-1). Inductors can be structurally closed with thermal insulation and protected by an outer screen from radiation and mechanical damage.

The use of a magnetically soft material in the core (by annealing electrical steel, for example 10860) can significantly reduce the hysteresis loop, increase the magnetic moment created by the electromagnet, while providing a low coercive force in the hysteresis loop, and therefore. the core practically does not remain magnetized when the field is removed, which is especially important when the direction of the magnetic moment is repeatedly switched (see, for example, A.G. Slivinskaya. Electromagnets and permanent magnets. M. Energia, 1972, pp. 22-24 ).

To improve the reliability of the magnetic system and the efficiency of the spacecraft control (increase the control torque), the electromagnets EM _X , EM _Y , EM _Z in the proposed invention are made of n electromagnets (n≥1), each of which contains two windings L1-i and L2- i. These n electromagnets form magnetic moments in one direction and can be connected in parallel or in series.

The relay node 11 used for switching the windings of the inductors in the block of electromagnets BEM 1 has a high reliability, in view of the fact that it is in standby mode, i.e. there is no main indicator of unreliability - a large number of switching contacts in the relay node 11.

It is known (see, A.P. Kovalenko. Magnetic control systems for spacecraft. M., "Mashinostroenie", 1975, pp. 169-170) that when current I flows in an inductor with a number of turns w and an average coil area S _cp , the coil creates a magnetic moment Q:

In addition, this source of information provides reasonable conclusions, from which it follows that, all other things being equal, the magnetic moment of the inductor will be the greater, the larger its dimensions.

The process of calculating the number of turns w, the wire diameter and the given current is iterative, because with each change in one parameter, it is necessary to recalculate everything, achieving the maximum ampere-turns, so the way to increase the magnetic moment Q (power flux) by increasing the ampere-turns is not rational. When designing electromagnets, an increase in the magnetic moment (power flux) is most optimal to use the conditions for reducing the magnetic resistance. To do this, it is necessary to choose a magnetic core with the shortest path of field lines and with the largest cross section, and as a material - an iron material with high magnetic permeability.

In this case, it is necessary to take into account the limitation of the maximum overall dimensions of electromagnets:

• overall dimensions and appearance of the spacecraft hull;

• optimal ratio of frame diameter and core length:

оптимальное соотношение - когда длина сердечника равна длине каркаса катушки;

the optimal ratio is when the length of the core is equal to the length of the coil frame;

длина сердечника не должна существенно превышать диаметр (предельное практическое значение максимального соотношения - приблизительно 10:1).

the length of the core should not substantially exceed the diameter (the limiting practical value of the maximum ratio is approximately 10:1).

Thus, taking into account a wide range of masses, dimensions and appearance of launched spacecraft, it becomes necessary to create a number of standard-sized electromagnets in EM _x , EM _y , EM _z , each of which has optimal parameters for operation in a magnetic system.

When currents flow in the coils of electromagnets EM _x , EM _y , EM _z of the block of electromagnets BEM 1 along the associated axes X, Y, Z of the spacecraft, components of magnetic moments Q _x , Q _y , Q _z are created, which, when interacting with the components of geomagnetic induction B _x , B _y , B _z of the geomagnetic field of the Earth (2), in projections on the associated axes of the spacecraft X, Y, Z, create control mechanical moments M _x , M _y , M _z (3), which are described (taking into account expression 1) with the following expressions:

When the spacecraft is separated from the rocket or upper stage, the process of calming is carried out, then the reduction process (combination of the axes of the associated spacecraft coordinate system with the axes of the orbital coordinate system) and subsequent stabilization of the spacecraft.

Practically all modern spacecraft use flywheel engines in the spacecraft orientation and stabilization system, which are distinguished by high stabilization accuracy, small size, ease of control, and the absence of non-renewable working fluid consumption. Flywheel motors create an internal control moment (they create a mechanical moment around an axis parallel to the axis of rotation of the flywheel motor rotor) and cannot fend off external disturbances, therefore, when using a block of flywheel motors 5 and a magnetic system that generates mechanical moments M _x , M _y , M _z , controlling the spacecraft body (4) is provided by:

• preliminary calming by suppressing the initial angular velocity of the spacecraft, acquired as a result of the separation of the spacecraft from the launch vehicle or upper stage, using only the influence of the control mechanical moments M _x , M _y , M _z (3) of the magnetic system. The equations of motion in the preliminary calming mode are given in the book by A.P. Kovalenko. Magnetic control systems for spacecraft. M., "Engineering", 1975, pp. 107-109;

• movement of the spacecraft around the center of mass (orientation and stabilization of the spacecraft) with the combined use of the magnetic system and flywheel motors of the DM 5 block;

• Resetting the kinetic momentum of the spacecraft in accordance with the law of conservation of angular momentum with the combined use of the magnetic system and motor-flywheels of the DM 5 unit.

As a flywheel motor in the DM 5 block, you can use, for example, DM 1-20 (or, for example, DM1-50), made on the basis of a controlled torque contactless DC motor, developed by VNIIEM Corporation JSC, Moscow ( see "Motors-flywheels for attitude control systems of spacecraft". On the website: https://www.vniiem.ru/) This flywheel motor DM1-20 has a pulse interface, an input for controlling the rotor speed and a built-in speed sensor rotor, through which the micro-computer 8 in a constant mode fixes the frequency of rotation of the rotor of the engine-flywheel in the block DM 5 and provides adjustment to achieve its required values.

To control the attitude of the spacecraft in block DM 5, at least three flywheel motors are required: for each of the orientation axes of the spacecraft (roll - rotation around the OXs axis, yaw - rotation around the OYs axis, pitch - rotation around the OZs axis).

It is not advisable to reserve them in the "cold" mode in view of the fact that, as practice has shown, an electromechanical device that has not been turned on in space for a long time has every chance not to turn on even at the moment when it is needed. Therefore, to increase reliability, it is more expedient to use a redundant system, for example, with a minimum redundant system, flywheel motors should be located on the faces of the pyramid, the control torque should be divided between all four flywheel motors and the failure of one of them should be parried by redistributing the torque over the remaining three.

The most important characteristic of a flywheel engine is its maximum kinetic moment that it can accumulate. The flywheel motors of the DM 5 block are "frozen" into the spacecraft body. those. does not make any movements relative to the body, but participate in its movement in accordance with the law of conservation of the angular momentum, which is described by the expression:

where J _a , J _m , respectively, the moments of inertia of the spacecraft and the flywheel:

ω _a , Ω _m respectively their current angular velocities;

ω _a (0), Ω _m (0) initial values of angular velocities

(See Scientific and Technical Bulletin of the St. Petersburg State University of Information Technologies, Mechanics and Optics, 2009, No. 5(63), p. 48).

The products J _a ⋅ω _а and J _m ⋅Ω _m are called momentum moments of the spacecraft and the flywheel, respectively.

From expression (5) it can be seen that any change in the angular velocity of rotation of the flywheel ΔΩ _m =Ω _m -Ω _m (0), leads to a change in the angular velocity of the spacecraft Δω _a = ω _a -ω _a (0), but in the opposite direction and in ratio determined by the moments of inertia.

Let us consider the process of formation by electromagnets of the components of magnetic moments Q _x , Q _y , Q _z .

The micro-computer 8 receives the components of geomagnetic induction In _x , By _y , B _z and components of the angular momentum of the spacecraft K _x , K _y , K _z .

Components of geomagnetic induction B _x , B _y , B _z are measured using a magnetometer 7.

As a magnetometer 7, you can use a three-dimensional magnetometer, for example, a digital three-component magnetometer MCT-11 developed by JSC "Ramenskoye instrument-making design bureau".

The components of the kinetic moment of the spacecraft K _x , K _y , K _z , in accordance with expression (5) correspond to the components of the kinetic moment of the flywheel engines in the DM 5 unit and are measured using angular velocity sensors BDUS 6, which is built on the basis of highly sensitive angular velocity sensors , measuring directly the angular velocity vector, and having the necessary resource. In order to increase the reliability and survivability of BDUS 6, with minimal hardware costs, it is advisable to use functional redundancy (four, six or more) based on the use of a non-orthogonal arrangement of angular velocity sensors. At the same time, due to the choice of a rational configuration of the angular velocity sensors, the accuracy of the BDUS 6 can be increased by 30-40% (see Parshin A.P., Nemshilov Yu.A. Development of a measuring unit of the UAV orientation system with a non-orthogonal arrangement of sensitive elements // Modern equipment and technologies, 2016, No. 3).

As angular velocity sensors in BDUS 6, for example, sensors based on microelectromechanical systems (MEMS) can be used. manufactured by the Microinstrument Laboratory Company, Zelenograd, or, for example, devices based on fiber-optic gyroscopes developed by CJSC NPP Antares BIUS-M-1.

According to the signals B _x , By _y , B _z and K _x , K _y , K _z entering the microcomputer 8, signals are generated at its information outputs equal to:

as well as the prohibition signal α _o , which, if it corresponds, for example, to a logical "0", prohibits the inclusion of the block of electromagnets BEM 1 and the formation of parasitic mechanical moments M _{x pairs} , M _{y pairs} , M _{z pairs} along the associated axes X, Y, Z of the space apparatus, and vice versa, if it corresponds to a logical "1", allows the inclusion of a block of electromagnets BEM 1 and the formation of useful mechanical moments M _{x field} , M _{y field} , M _{z field} along the associated axes X, Y, Z of the spacecraft. This is necessary so that in the case when the parasitic moment is greater than the useful one, the device can not only not reset the kinetic moment, but, on the contrary, increase it (parasitic disturbances when the magnetic actuators are turned on arise due to the fact that the mechanical moment created by the interaction of the magnetic field electromagnets with a geomagnetic field, in the general case, is the sum of two components - a useful component, the vector of which is parallel to the angular momentum vector and directed in the opposite direction, and a parasitic component, the vector of which is perpendicular to the angular momentum vector).

This can be seen in the example of the formation of useful and mechanical magnetic moments along the associated X axis from the graphs presented in Fig. 2 (according to expressions (6) in the formation of mechanical moments along the associated axes Y and Z, the processes are similar):

находится под произвольно выбранным углом к вектору

в пределах от 0 до 90°;• in FIG. 2(a) vector

is at an arbitrary angle to the vector

within the range from 0 to 90°;

находится на оси вектора

• in FIG. 2(b) vector

is on the axis of the vector

находится под углом 90° к вектору

• in FIG. 2(c)vector

is at an angle of 90° to the vector

перпендикулярной вектору

. Эта составляющая геомагнитной индукции при включении магнитного исполнительного органа, формирующего магнитный момент L_x, создает полезный механический момент

направленный в противоположном направлении к вектору

. Величина B_YZCos_γ равна проекции вектора

на направление вектора

. При включении магнитного исполнительного органа эта составляющая геомагнитной индукции является причиной возникновения паразитного механического момента

направление которого перпендикулярно составляющей

.From the graphs in Fig. 2 (a) it can be seen that the factor B _YZ Sinγ is equal to the value of the component of the vector

perpendicular to the vector

. This component of geomagnetic induction, when the magnetic actuator is turned on, which forms the magnetic moment L _x , creates a useful mechanical moment

directed in the opposite direction to the vector

. The value B _YZ Cos _γ is equal to the projection of the vector

to the direction of the vector

. When the magnetic actuator is turned on, this component of geomagnetic induction is the cause of the occurrence of a parasitic mechanical moment

whose direction is perpendicular to the component

.

и

полезный механический момент

отсутствует, а паразитный

имеет максимальное значение.From the graphs in Fig. 2 (b) it can be seen that if the directions of the vectors coincide

And

useful mechanical moment

absent, parasitic

has the maximum value.

и

полезный механический момент

имеет максимальное значение, а паразитный

отсутствует.From the graphs in Fig. 2 (c) it can be seen that with the perpendicular direction of the vectors

And

useful mechanical moment

has a maximum value, and parasitic

absent.

что соответствует значению угла γ>45° и наоборот сигнал запрета a _о, соответствующий логическому "0", формируется при условии

что соответствует значению угла γ<45°.As can be seen from the graphs in Fig. 2 prohibition signal a _o , corresponding to logical "1", is generated under the condition

which corresponds to the value of the angle γ>45° and vice versa, the prohibition signal a _about , corresponding to the logical "0", is formed under the condition

which corresponds to the value of the angle γ<45°.

випуск 4(14), 2001). Так, например, в качестве микро-ЭВМ 8 возможно применение 4-х канальной ЦВМ201 (разработка АО «НИИ «Субмикрон», г. Зеленоград), при этом, рекомендуемым вариантом является мажоритированная работа 3-х модулей, а четвертый находится в холодном резерве.To improve the reliability and survivability of microcomputers 8, it is advisable to perform it in the form of a majorized three-channel structure (see, for example, N.K. Baida et al.

issue 4(14), 2001). So, for example, as a micro-computer 8, it is possible to use a 4-channel TsVM201 (development of JSC "Scientific Research Institute" Submikron ", Zelenograd), while the recommended option is the majority operation of 3 modules, and the fourth is in cold reserve .

Let us consider the process of formation of magnetic moments Q _x , Q _y , Q _z , respectively, by electromagnets EM _x , EM _y , EM _z in the block of electromagnets 1.

As was accepted in the description above, the formation of magnetic moments Q _x , Q _y , Q _z is carried out with a prohibition signal a _o corresponding to a logical "1", which is formed in the micro-computer 8 and fed to the input of the DC voltage converter unit 15, namely , to the "prohibition-enable" inputs of the microcontrollers MK 16 _i DC voltage converters PPN _i 15 _i (where i means x, y, z), and "allows" the formation of PWM signals (pulse-width modulated pulse sequence), in accordance with the signals σ _x , σ _y , σ _z , coming from the group of information outputs of the microcomputer 8. The PWM signals generated in the DC voltage converters PPN _i 15 _i form alternating pulse voltages at the outputs of the inverters AND 17 _i with a constant amplitude and variable pulse duration, which are converted into constant voltage at the outputs of the rectifiers 18 _i . The generated DC voltage from the outputs of the DC/DC converter unit 15 is supplied to the reversible switches 9 _i in the form of power supply voltages (+U _pit , -U _pit ) _i .

The DC voltage converters PPN _i 15 _i are powered by the onboard power supply of the spacecraft UPS KA 19. For example, the rated onboard DC power supply of the Kanopus-V spacecraft is 27 V.

DC voltage converters PPN _i 15 _i allow you to generate output voltages corresponding to the signals σ _x , σ _y , σ _z , with high accuracy by changing the duty cycles D _i in a pulse sequence with PWM in accordance with changes in the signals σ _x , σ _y , σ _z .

The fill factor D is described by the following expression:

Where

τ _imp - pulse duration in a pulse sequence with PWM;

T - the period of repetition of pulses in a pulse sequence with PWM.

Inverters And 17 _i in DC voltage converters PPN _i 15 _i can be made in the form of bridge converters with PWM control (see, for example, PWM bridge converter. On the website: https://power-electronics.info/full-bridge. html). Bridge circuits of inverters And 17 _i , made on the basis of a single-phase inverter, are a high frequency link (typical high frequencies are above 20 kHz. In fact, in the range from 40 to 100 kHz).

These pulse converters have a high efficiency q, which reaches more than 90% (see, for example, "How pulse voltage converters work". On the website: https://www.qrz.ru/schemes/contribute/power/kak- rabotajt-impul-snye-preobrazovateli-naprazenia-27-shem.html).

Rectifiers 18 _i can be made in the form of a diode bridge with an output filter (see, for example, "Diode Bridge". On the website:

https://www.ruselectronic.com/diodnyj-most/).

Microcontrollers 16 _i can be used, for example, microcontrollers ST10F276Z5T3.

Let us consider the formation of currents in electromagnets in one channel, for example, a channel along the associated X axis of the spacecraft, since the operation of the channels along the associated axes Y, Z is similar, in view of the fact that the channels X, Y, Z of the block of electromagnets 1, relay node 11, DBK 9 and current sensors 10 are made according to an identical scheme.

If the signal a _about logical "1" corresponds, the control device for reversing switches UU RK 12 opens one of the diagonals of the transistor bridge of the bridge circuit 9 _x , which determines the direction of the current in the block of electromagnets BEM 1.

For a given direction of current in the inductors L1-j, L2-j (where j means 1, 2, ... n) of the electromagnet EM _x (for example, provided by open transistors V1 and V4 of the 9 _x bridge circuit), the current flows through the circuit: positive power bus (+ U _pit ) - transistor V1 - relay node 11 - inductors L1-j, L1-j - relay node 11 - LED of the optocoupler node of the current sensor 10 _x -2 - transistor V4 - negative power bus (-U _pit ). At the same time, at the output of the photodetector (F) of the optocoupler assembly of the current sensor 10 _x -2, an electrical signal is generated, which, entering the micro-computer 8 as a feedback signal, "confirms" the given (we take for direct) direction of the current of the inductors L1- j, L2-j of the electromagnet EM _x and, accordingly, the direction of the magnetic moment created by the electromagnet EM _x , which, when interacting with the geomagnetic field of the Earth, creates a vector of the control mechanical moment on the spacecraft. If the direction is not "confirmed" by the feedback signal, then the micro-computer eliminates this error at the input of the control device for reversing switches UU RK 12.

As transistors in bridge circuits, it is advisable to use transistor modules in the form of two transistors included in the diagonal of the bridge. In the block of reversing switches BRK 9, as well as in inverters AND 17 _i , you can use modules based on modern insulated gate field effect transistors (MOSFETs), which have the best dynamic characteristics in comparison with other types of transistors, for example, IGBT insulated gate bipolar transistors (see Fig. , patent, RF, No. 2426215).

When changing the direction of the current in the inductors L1-j, L2-j, at the signal of the micro-computer 8, the control device for reversing switches UU RK 12 opens another diagonal of the transistor bridge of the bridge circuit 9 _x (transistor V3 and V2 opens) and the current flows in the opposite direction along the circuit: positive power bus - transistor V3 - LED of the optocoupler node of the current sensor 10 _x -1 - relay node 11 - inductors L1-j, L2-j - relay node 11 - transistor V2 - negative power bus. At the same time, already at the output of the photodetector (F) of the optocoupler unit of the current sensor 10 _x -1, an electrical signal is formed, which, entering the micro-computer 8 as a feedback signal, "confirms" the given (opposite) direction of the current of the inductors L1-j, L2-j electromagnet EM _x that creates a given (opposite) direction of the magnetic moment and, accordingly, the direction of the control mechanical moment.

Separate electromagnets, consisting of coils (L1, L2)-j, operating in hot standby mode (connected in parallel), are connected in series or in parallel in the electromagnet EM _x , to which, from the block of reversible switches BRK 9 through the relay node 11, a constant voltage U is supplied _el corresponding with high accuracy to the signals σ _x , σ _y , σ _z .

Taking into account Ohm's law and the dependence of current on voltage in inductance (see, for example, Matvienko V.A. Fundamentals of circuit theory. Textbook. Ekaterinburg, 2016. S. 32, 35)) processes occurring in electromagnets EM _x , EM _y , EM _z are described by the expression:

where I is the current in the coil, which increases exponentially when voltage is applied to the electromagnet coil;

R - ohmic (active) resistance of the inductor, causing irretrievable losses of DC energy;

E _equiv is the equivalent inductance that forms the magnetic moment at a given moment of time by the inductances of the electromagnets EM _x , EM _y , EM _z of the BEM 1 block.

The first term (I•R) in expression (8) characterizes the amount of heat release by electromagnets EM _x , EM _y , EM _z , and the second - the value of the magnetic moment created by electromagnets EM _x , EM _y , EM _z .

By connecting inductors in series or in parallel, an increase or decrease in the total (equivalent) inductance is provided, and thus the required optimal equivalent inductance can be obtained. Known (see, for example, "Parallel, series connection of chokes".

On the website: https://gyrator.ru/paraHei-serial-inductance) that the equivalent inductance L _equiv when n coils are connected in series with inductance L is generally described by the expression:

and with a parallel connection, it is described by the expression:

where L ₁ , L ₂ ... L _n in expressions (9, 10) - respectively, the inductance of the "working" coils (L1, L2)-j in electromagnets EM _x , EM _y , EM _z .

When connecting coils (L1, L2)-j in series or in parallel in electromagnets EM _x , EM _y , EM _z of the BEM 1 unit, it is necessary to comply with the condition of the same direction of currents in them (respectively, the directions of magnetic moments formed), as well as the optimal choice from a number of standard sizes electromagnets for pairing them with the onboard power supply of the spacecraft in terms of voltage and maximum current. As a result, the BEM 1 block will generate high magnetic moments along the associated X, Y, Z axes of the spacecraft, and, accordingly, the effective total mechanical effect of individual n electromagnets on the spacecraft body and its control.

The electrical signals entering the microcomputer 8 from the output of the photodetectors (F) of the optocoupler units of the current sensors 10 _x -1 and 10 _x -2 can also be used as telemetry signals (TM) to ensure the appropriate switching of the inductors L1-j, L2 -j electromagnet EM _x and/or blocks of the main and backup channels DBK 9 and current sensors 10 (not shown in Fig. 1. The design of the main and backup channels of the reversible switch with current sensors is shown in the prototype).

As a photodetector (F) in optocoupler units of current sensors 10 _x -1 and 10 _x -2, there can be either a photodiode, or a phototransistor, or a photoresistor (for example, see Electronic technology in automation, issue 10, Moscow, 1978, p. 210.).

Protection against through short-circuit currents when switching power transistors of the bridge circuit BRK 9 is given, for example, in the description of the inductive load reversing switch in the patent, RF, 2140128.

Breaks and interturn short circuits in the inductors L1-j, L2-j of individual electromagnets, as well as violations in the connections between individual electromagnets, are accompanied by changes in the current levels in the current sensor unit 10, as well as changes in the levels of magnetic moments created by the electromagnets EM _x , EM _y , EM _z of the block of electromagnets BEM 1 and, consequently, the control mechanical moments on the spacecraft. In this case, the micro-computer 8 (if necessary, using signals from the NKU 14) controls the block of relay switches 9 (first of all, voltages ± U _pit ) and the relay node 11 for the appropriate choice of connections between the coils in the block of electromagnets BEM 1 and the normalization of modes work of inductors L1-j, L2-j.

Some individual turn-to-turn short circuits in the inductors L1-j, Lk2-j can lead during operation only to a slight decrease in the inductance of the electromagnets, which does not significantly impair the efficiency of the magnetic system. To improve the reliability of the magnetic system, interturn short circuits in the inductors L1-j, L2-j of the electromagnets EM _x , EM _y , EM _{x of} the electromagnet unit BEM 1, it is advisable to detect on special stands before installing them on the spacecraft by passing alternating current of a certain frequency through them .

It should be noted that for large spacecraft that use two-stage flywheel motors (gyrodynes) with a large control torque, it is advisable to use the combined use of a magnetic system with an attitude control jet unit with an insignificant supply of the working fluid (see, for example, patent, RF No. 2767648) .

As a result of this combined use, the time of the process of calming and stabilizing the spacecraft is significantly reduced, "parasitic" angular velocities of the spacecraft are effectively parried during the spin-up of the gyrodyn rotors, and conditions are also provided for optimizing the choice of batteries and solar batteries, which make it possible to increase the specific energy characteristics of the onboard power source of the spacecraft UPS CA 19.

Claims (1)

A device for resetting the angular momentum and controlling the attitude of a spacecraft using a magnetic system, which includes an electromagnet with an inductance coil consisting of two inductors - the main and backup coils, a micro-computer, a reversible switch made according to a bridge circuit on semiconductor power elements, and a device control of reversing switches, the output is connected to the control inputs of the reversing switch, one diagonal point of which is connected through a relay node to one end of the main inductance coil of the electromagnet, the second end of which is also connected through the relay node to one output of two parallel oppositely connected LEDs of two optocoupler nodes of current sensors, the second output of which is connected to the second point of the diagonal of the reverse switch, while a separate output of the micro-computer is connected to the input of the reverse channel control device, and separate inputs are connected to the outputs of the photodetector of the optocoupler units of the current sensors, characterized in that an additional block of electromagnets, a block of reversible switches, a block of current sensors, a block of DC voltage converters, a block of angular velocity sensors, a block of flywheel motors and a magnetometer, while the block of electromagnets contains electromagnets installed along the associated X, Y, Z axes of the spacecraft, respectively, electromagnets EM _x , EM _y , EM _z , each of which consists of n (n≥1) electromagnets connected in series or in parallel through a relay assembly with a block of reversible switches, and containing the main and backup inductors operating in hot mode; block of reversible switches, consisting of three identical reversing switches along the connected axes X, Y, Z of the spacecraft, each of which is connected to the outputs of the control device for the reversing switches and the outputs of the block of DC voltage converters, while each reversing switch is connected to identical devices of the optocoupler units of the block current sensors, respectively, along the associated axes X, Y, Z of the spacecraft, which are connected to the relay node and a group of separate inputs of the microcomputer; the block of angular velocity sensors, the block of flywheel motors and the magnetometer are connected to a group of separate inputs of the micro-computer, which is connected by a group of information outputs to the input of the control device for reversing switches, the inputs of the DC voltage converter unit and to the microcontrollers in each DC voltage converter along the connected X axes, Y, Z of the spacecraft, while the PWM outputs of the microcontrollers are connected to the control inputs of the inverters, the outputs of which are the outputs of the DC voltage converter unit through rectifiers, and are connected via the power buses to the onboard power supply of the spacecraft.

Images