RU2618664C1 - Method of spacecraft orientation and device for the method realisation - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: group of inventions refers to the methods and means of satellites magnetic orientation, mainly of small spacecrafts (SC). The method includes measuring the Earth's magnetic field induction vectors and kinetic moment, accumulated by the "KA-flywheel" system. The signals to the executive electromagnetic devices for the preliminary spacecraft calming are formed according to the measured parameters. The gyrostabilization and the spacecraft motion control are carried out in the pitch with the help of the engine-flywheel (installed along the pitch axis). The flywheel unloading is carried out by means of the indicated electromagnetic devices. These processes are controlled by the microcomputer, interacting with the on-board computer, connected through the telemetry channel to the ground center.

EFFECT: improving of the technical and operational characteristics and the orientation system reliability, mainly of small spacecrafts.

2 cl, 1 dwg

Description

Appointment

The invention relates to space technology, in particular to means of orientation of small spacecraft (SC), moving in orbit in a magnetic field and equipped with an orientation system using the Earth's magnetic field.

State of the art

The most important task of modern space technology is the creation of small-sized spacecraft. Currently, small spacecraft, the mass of which is less than 100 kg, are in demand. The reasons for the rapid development of this class of satellites are the relatively low cost and the insignificant time required for the design and manufacture of a small spacecraft, the ability to launch alongside at a low cost of launching into orbit. The class of tasks performed with the help of small spacecraft is very wide. For example, microsatellites can be used to capture the Earth in detailed mode with a resolution of 6-10 m, as well as in a wide band with a resolution of hundreds of meters to 2 km. Such satellites can solve fire detection tasks, conduct surveys of natural disaster zones, conduct environmental monitoring, and can be used to develop new technologies and conduct experiments in space. Of particular relevance is their use for scientific research and remote sensing of the earth, where the development of scientific and technical solutions and methods for creating small (ultra-small) spacecraft of a multi-tiered monitoring system, for example, the radiation situation in near-Earth space and determining their shape, is required (see Boyarchuk K. A., Karelin A.V., Makridenko L.A. “Prospects for monitoring radioactive contaminants from outer space on the Earth's surface and in the lower atmosphere”. Electromechanics. - 2005. - T. 102. - P. 183-209).

For example, a multi-tiered grouping may consist of 3 (4) spacecraft operating in orbits:

- one spacecraft in orbit with an altitude of 650-700 km with an inclination of 85-95 °;

- one (two) spacecraft in orbit with an altitude of 1500-2000 km with an inclination of 65-70 ° (110-115 °);

- one spacecraft in orbit at an altitude of 7500-8000 km with an inclination of 45-50 ° (130-135 °).

Along with the requirements for the accuracy of orientation and stabilization of small spacecraft with the simultaneous manufacture of several small spacecraft, high technical and operational requirements are put forward:

- minimum weight, dimensions, power consumption and cost;

- high reliability (the life of the MCA is at least 5 years);

- maximum simplicity;

- high technology;

- minimum time for ground mining, testing and preparation for launch.

Analyzing alternative options in relation to solving the problem, we consider the most famous system - a system with gas-jet nozzles or micro-jet engines. These executive bodies, throwing cold or hot gas through a nozzle into outer space, create a force acting on the device. Since the reserves of working gas on board the spacecraft are limited and not replenished, for a long flight for small spacecraft this system is unacceptable because of the required large reserves of working gas, leading to a significant increase in mass and dimensions.

For systems using gravitational fields, aerodynamic pressure, solar radiation forces, only the stabilization function of the spacecraft in one basic position is characteristic: in the local vertical of gravitational systems; in the direction of the flight velocity vector for aerodynamic systems and, finally, in the direction to the Sun for systems with “solar sails”. These systems, although they have the fundamental ability to perform other functions (preliminary reassurance, spatial turns, etc.), are nevertheless never used for this in view of their apparent inappropriateness due to their excessive complexity (see Gushchin V.N. Orientation and stabilization systems // Fundamentals of spacecraft: Textbook for high schools. - M .: Mashinostroenie, 2003. - S. 241-257. - 272 p.).

At present, the flywheel magnetic orientation system, which is one of the main spacecraft onboard systems that provides a certain position of the spacecraft's hull in space, is widely used. As the executive bodies, one of the main elements of this orientation system are flywheel engines. With its help, control moments are generated to obtain the required spacecraft orientation. Currently known electromechanical actuators based on flywheel engines consist of 3 engines mounted on a rigid base. Their rotation axes form an orthogonal coordinate system and coincide with the main axial moments of inertia of the spacecraft (see, for example, A. G. Iosifyan. Electromechanics in space. - “Cosmonautics, Astronomy”, 1977, No. 3, p. 20).

The flywheel engine is an electric motor with a massive flywheel mounted on its shaft. However, systems in which flywheel engines are used as executive bodies cannot control the magnitude of the total kinetic moment of the inertial mechanical system “SC body - flywheels”, since they are not able to create moments external to the SC.

For these purposes, magnetic systems that are not much inferior in terms of functionality to widely known reactive systems (except that they cannot create a control moment relative to the direction coinciding with the Earth's magnetic field induction vector, and independent moments are simultaneously relatively suitable for spacecraft) all control axes). Compared to them, they have undeniable advantages in that, for example, electromagnets can be used as magnetic actuators. They do not consume fuel or gas at all, i.e. working fluid (see AP Kovalenko. Magnetic control systems for spacecraft. M., "Mechanical Engineering", 1975, p. 9).

Flywheel engines are most widely used to control the orientation of small spacecraft, the mass of which is less than 500 kg. It should be noted that the untwisted engine-flywheel with a constant kinetic moment can be used as a gyrostabilizer. It is known that a gyroscope has a very important property to maintain the direction of its main axis unchanged in inertial space (see V. A. Pavlov. Gyroscopic effect of its manifestation and use, L., Shipbuilding, 1985, p. 6).

A known method of using the gyroscopic moment for controlling an aircraft (vehicle) and an aircraft control device (RF patent No. 2495789).

A method of controlling an aircraft, i.e. changes in its orientation in space and / or its stabilization consists in using the gyroscopic moment created by the gyroscope without the need for “unloading” gyroscopes, as well as using gyroscopes having a small mass relative to the aircraft itself in comparison with other orientation and stabilization systems using gyroscopic moment.

The control device of the aircraft, the operation of which is carried out using the specified method, consisting of a housing in which a central axis with rods is fixed through a system of additional torque, two gyroscopes having identical sizes and weights are attached to each rod.

The rotation vectors of the first and second gyroscopes are opposite to each other. The moment vectors of the first and second gyroscopes also have opposite directions and neutralize the reverse actions of each other. Accordingly, the action of the moment of one gyroscope on another gives rise to the appearance of Coriolis forces that create the moments of each gyroscope. The sum of these moments creates the moment that rotates the central axis of the device. Thus, if the central axis of the device is rigidly connected to the spacecraft, then under the influence of this moment, it will turn in the direction of this moment.

The invention eliminates the need for "unloading" of the gyroscope, however, there are two gyroscopes in the device, with the requirement of identity in size and weight, as well as additional electromechanical devices (additional weight and dimensions). Taking into account the control along the three orthogonal axes X, Y, Z (the main axial moments of inertia of the spacecraft), gyroscopes located along these orthogonal axes are necessary (see ER Rakhteenko. Gyroscopic orientation systems. M., "Engineering", 1989 , p. 10-12).

Therefore, the use of this system, even in the absence of duplication, for small (ultra-small) spacecraft with the technical and operational requirements put forward to them, described above, is not a rational solution.

This disadvantage can be eliminated by using the most suitable magnetic systems for these purposes, which, as mentioned above, are practically inferior in functionality to reactive systems and do not consume a working fluid at all.

This principle of orientation and stabilization is carried out in the orientation method and the orientation system for implementing the method according to the patent of the Russian Federation No. 2150412, taken as a prototype.

This method of orienting the spacecraft consists in creating an ellipsoid of inertia orienting the spacecraft in the Earth's gravitational field in the orbital coordinate system, by moving the structural elements of the spacecraft at predetermined distances relative to each other, in the primary measurement at fixed time intervals of signals proportional to the projections of the magnetic field of the Earth axis associated with the spacecraft coordinate system, storing the magnitude and polarity of the measured signals until the next primary measurement, creating a damping of its magnetic moment along each axis of the spacecraft, secondary measurements of signals are introduced, which are proportional to the projections of the Earth's magnetic field induction vector, continuously or with time intervals shorter or equal to the memory intervals of primary measurements of signals, and secondary signals are compared on each axis with the corresponding stored primary signals measurements, receive comparison signals and create continuously or at time intervals between secondary measurements a damping magnetic moment for each Axis of the spacecraft, the value of which is constant or proportional to the magnitude of the corresponding signal comparison spacecraft axis, and the direction opposite to the polarity of the signal.

The orientation system that implements this method contains a gravity device and three damping channels into which comparison blocks are introduced. These blocks are associated with the outputs of the magnetometers, memory blocks and devices for creating magnetic moments of the spacecraft.

Disadvantages that limit the use of this patent in small spacecraft:

- sufficiently large mass and dimensions of the gravitational device and an increase in the maximum moment and the steepness of the moment characteristics in this system leads to an increase in mass and dimensions;

- low steepness of the moment characteristics of the gravitational system, providing low accuracy at disturbing moments;

- unique complex mechanical devices require the use of expensive technological processes and lead to a decrease in the reliability of the spacecraft orientation system.

In addition, the gravitational stabilization device has such operational disadvantages as, for example, the deformation of the rods (or other mechanical parts) from uneven thermal heating by the sun's rays (see RF patent No. 2304069), the elimination of which requires additional original solutions that complicate expensive technological processes and increasing the dimensions, mass and cost of the spacecraft, and leading to a decrease in the reliability of the operation of the spacecraft orientation system.

The aim of the invention is to increase the reliability of the orientation system of the spacecraft.

Disclosure of invention

This goal is achieved due to the technical support of the required technical and operational characteristics of small spacecraft moving in the Earth’s near-Earth orbit, oriented in the orbital coordinate system, by using a magnetic orientation system in conjunction with one flywheel engine used as a gyro stabilizer. A reliable magnetic system, which is close in functionality to reactive systems, does not consume a working fluid, provides preliminary calming of the spacecraft and, subsequently, control along the coordinate axes (roll, yaw, pitch). Used only one flywheel engine with a kinetic moment directed along the binormal to the plane of the orbit (along the spacecraft axis), as a gyrostabilizer and pitch control to insignificant limits, determines the overall high reliability of the small spacecraft orientation system.

The proposed method of orienting the spacecraft is to determine the magnitude of the magnetic induction of the Earth’s magnetic field along the three coordinate axes X, Y, Z and the accumulated kinetic moments of the inertial mechanical system “spacecraft-flywheel body” from roll, yaw, pitch, formation of magnetic moments at the outputs of electromagnetic devices, which, interacting with the Earth’s magnetic field, create controlling mechanical moments for pre-calming and controlling the spacecraft. In addition, in gyro stabilization and spacecraft pitch control using a flywheel engine, determining the accumulated kinetic moment of a flywheel engine, and forming magnetic moments at the outputs of electromagnetic devices that interact with the Earth’s magnetic field, create a controlling mechanical moment to reset the kinetic moment of the engine flywheel.

The spacecraft orientation device includes three magnetometer sensors, the outputs of which are connected to the inputs of an electronic computing device, the outputs of which are connected to the inputs of three amplification-conversion devices, the outputs of which are connected to the inputs of three electromagnetic devices.

The introduction of three sensors of the kinetic moment component, the outputs of which are connected to the inputs of the electronic computing device, the flywheel engine connected to the output of the electronic computing device through the interface device, into the device for orienting the spacecraft, allows using magnetic devices to control magnetic moments along three coordinate axes X, Y, Z, which interacting with the induction vectors of the Earth's magnetic field create control moments for the spacecraft.

In the general case, the control moment M is determined by the basic control equation

where L is the vector of the generated magnetic moment by electromagnetic devices;

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

In projections and control axes of the spacecraft, expression (1) takes the form

(see AP Kovalenko. Magnetic control systems for spacecraft. M., "Mechanical Engineering", 1975, pp. 21-22).

The first feature of the spacecraft’s magnetic control system, as follows from formula (1), is that the vector of the control moment M is perpendicular to the vector B and therefore it is impossible to create a control moment in the direction of the field, i.e. all possible positions of M are enclosed in a plane normal to B. Therefore, in non-optimal sections of the spacecraft’s flight in orbit (control moment in the field direction), the influence of disturbances on the spacecraft cannot be eliminated and its control accuracy is low. This drawback of the magnetic control system, namely, its inability to create a control moment in the direction of the Earth's magnetic field induction vector, is “compensated” by introducing into the device a flywheel engine operating in the gyrostabilizer mode and allowing to weaken the influence on the orientation accuracy of this drawback. The installation of a gyrostabilizer on a spacecraft gives the latter the ability to resist, like a gyroscope, the action of a disturbing moment. This property turns out to be especially useful in those parts of the orbit in which the induction vectors of the geomagnetic field and the disturbing moment turn out to be almost parallel. Moreover, by changing the speed of rotation of the flywheel within small limits relative to the nominal speed of rotation, an additional control moment is created that allows you to adjust the motion of the spacecraft (pitch). The choice of the gyrostabilizer location by pitch is related to the fact that the spacecraft in orbital flight, oriented in the orbital coordinate system, rotates relative to the pitch axis. With this arrangement of the gyrostabilizer, it does not interfere with such a rotation of the spacecraft (does not create a gyroscopic moment). At the same time, when disturbing moments appear, the projection of the velocity perpendicular to the axis of rotation of the gyrostabilizer, it creates a gyroscopic moment that prevents the direction of the velocity vector from changing.

Using the flywheel control engine, by changing the speed of the flywheel in small limits relative to its rated speed, the required control moment along the pitch axis is maintained.

The combination of basic control with the help of magnetic moments created by electromagnets with flywheel motor control provides complete controllability along the pitch axis. On the roll and yaw axis, the system is controllable on average over the orbital period. This is explained by the fact that during the spacecraft’s flight in orbit B changes both in magnitude and direction relative to the control axes, and new conditions for control are created each time. If at the moment these conditions are in some sense not optimal, then after a while they will be close to optimal. The stabilizing properties of the gyrostabilizer during the orbital period keep the angles of deviation of the spacecraft within acceptable limits.

It should be noted that with the help of controlling magnetic moments along the three coordinate axes X, Y, Z created by electromagnetic devices and a three-component KA kinetic moment sensor, it is easy to provide both pre-calming of the KA and unloading the kinetic moment of inertial actuators (see B. I. Boyevkin et al. Orientation of artificial satellites in gravitational and magnetic fields. Nauka, Moscow, 1976, p. 211).

Given the development trend of microelectronics and software, miniaturization of electronics, the transition to digital technologies in control, the most optimal is the use of microcomputers as an electronic computing device. This is facilitated by the requirements of GOST R 52070-2003 on the use of standardized interfaces of serial multiplexed communication channels (MCO), which include an on-board computer with system and local controllers.

Thus, due to the introduction of new features - a three-component kinetic moment sensor, together with electromagnetic devices, the spacecraft is pre-calmed down and the kinetic moment of the inertial mechanical system is reset, and the introduction of the flywheel engine operating in the gyrostabilizer mode significantly reduces the influence of the disturbing moment on the spacecraft, more high orientation accuracy and the specified angular motion of the spacecraft hull is supported.

In addition, the proposed device provides comprehensive telemetry and microprocessor control of the device’s operating modes with high reliability, the specified orientation accuracy is achieved using a reliable flywheel engine and a reliable active magnetic orientation system, which has maximum simplicity and minimal mass inertia parameters, energy consumption and cost, - In general, this determines the high reliability of the onboard equipment of the orientation system.

Graphic illustration

FIG. 1. The structural diagram of the device orientation of the spacecraft.

Description of the spacecraft orientation method

The magnitude of the magnetic induction of the Earth’s magnetic field is determined by the three coordinate axes X, Y, Z and the accumulated kinetic moments by the inertial mechanical system “KA-flywheel body” by roll, yaw, pitch, form control electric signals for electromagnetic devices on each coordinate axis X, Y , Z, at the outputs of which magnetic moments are formed, which, when interacting with a geomagnetic field, create controlling mechanical moments for pre-calming and controlling the spacecraft, maintain a constant the rotational speed of the rotor of the flywheel engine to ensure the role of the spacecraft gyrostabilizer, further regulate the speed of the flywheel engine by controlling the rotational speed of the rotor using a microcomputer, to maintain the necessary angular velocity of the spacecraft in pitch, determine the kinetic moment of the flywheel engine and reset it using electromagnetic devices.

Device Description

The device for orienting a spacecraft using the earth’s magnetic field contains the components indicated by the positions in FIG. one:

- sensors of components of geomagnetic induction along the three coordinate axes X, Y, Z (M _x , M _y , M _z ) - 1, 2, 3;

- microcomputers (electronic computing device) - 4;

- amplifier-converting devices along the three coordinate axes X, Y, Z (UPU _x , UPU _y , UPU _z ) - 5, 6, 7;

- electromagnetic devices along the coordinate axes X, Y, Z (EM _x , EM _y , EM _z ) - 8, 9, 10;

- sensors of the components of the kinetic moment (DKKM) roll, yaw, pitch - 11, 12, 13;

- engine-flywheel (DM) - 14;

- interface device (CSS) - 15;

- IIM (information interface module) - 16.

When the spacecraft is separated from the rocket or the booster block, the process of calming down and the subsequent stabilization of the spacecraft in the orbital coordinate system OX _O Y _O Z _O :

- the origin coincides with the center of mass of the MCA;

- the axis Y _{O is} directed along the radius vector of the orbit of the MCA from the center of the Earth;

- the X _O axis is directed towards the linear velocity vector of the MCA, is in the orbit plane and is orthogonal to the Y _O axis;

- the axis Z _O complements the coordinate system to the right orthogonal.

Initially, the initial damping mode is automatically started, designed to reduce the kinetic moment (quenching of the angular velocity of rotation) of the spacecraft that occurs during separation and to stabilize the spacecraft with respect to an arbitrary inertial coordinate system determined at the time of separation, i.e., zeroing the initial kinetic moment of the spacecraft.

This is as follows.

The sensors of the components of geomagnetic induction M _x , M _y , M _z - 1, 2, 3 generate the signals of the components of the magnetic induction B _x , B _y , B _z , and the sensors of the components of the kinetic moment by roll, yaw, pitch - 11, 12, 13 measure kinetic moments of the spacecraft and generate signals K _x , K _y , K _z .

As sensors, components of geomagnetic induction can be used, for example, a three-component digital magnetometer MTTs-8 developed by Ramenskoye Instrument-Making Design Bureau JSC.

The angular velocity sensors, for example, sensors based on microelectromechanical systems (MEMS) manufactured by the Laboratory of Microdevices ”can be used as sensors for the components of the kinetic moment in roll, yaw, and pitch — 11, 12, 13 (see RF patent No. 2281232). , Zelenograd.

Based on the incoming signals, the microcomputer 4 generates signals equal to

Microcomputer 4 includes the following integral elements: microprocessor, RAM, ROM, data input and output devices (see RF patent No. 106768).

The microprocessor performs all the functions of organizing the operation of the microcomputer 4, synchronizing the interaction of its components and performing the necessary computational operations, selects instructions and data from the ROM program memory and data memory, performs arithmetic and logical operations on the data, and controls the signals on the internal data address bus. A microprocessor is the main program-controlled element that implements the process of processing digital information. It generates address signals and control pulses necessary to access the memory and input / output device, and also responds to interrupt requests.

In accordance with the software embedded in the ROM, the microprocessor organizes the work in several hard cycles.

When the microprocessor accesses the RAM, information is written to the RAM memory cells or information is transferred from the RAM memory cells; in the absence of circulation, the RAM cells are in the information storage mode.

As a microcomputer, you can use, for example, domestic single-chip microcomputers of the 1816 series. The use of single-chip microcomputers ensures the achievement of exceptionally high performance indicators (all computational algorithms for processing input information and generating the required controls for electromagnetic devices, a flywheel engine are implemented) at a low cost.

Command and program information (KPI) and navigation data (ND) to the microcomputer 4 is received through the information interface module 16 (taking into account telemetry data, for example, according to GOST R 52070-2003). The main bus of the information transmission line is made of a cable with a twisted shielded pair of wires in a protective sheath, to which matching elements are connected from both ends of the cable. As the information interface module 16, for example, a USB TA1-USB interface module with a multiplex channel according to GOST R 52070-2003 (MIL-STD-1553 V) can be used. This information interface module is capable of processing about 500 messages per second. In the interfaces of multiplex communication channels, information is exchanged according to the command-response principle with time division of messages consisting of command, information and response words. Control commands from a local controller (on-board control complex), as well as telemetry information (via a central computer) on the operation of all key elements of the device and microcomputer 4 (and vice versa, read information into a central computer that is telemetry (TM), are transmitted via a multiplexed communication channel) connected to the ground control center).

The received signals σ _x , σ _y , σ _z from the computer 4 data output device are fed to the inputs of the amplification-conversion devices UPU _x , UPU _y , UPU _z - 5, 6, 7, which provide the formation of currents in the coils of electromagnetic devices EM _x , EM _y , EM _z - 8, 9, 10. As a result, at the outputs of electromagnetic devices EM _x , EM _y , EM _z - 8, 9, 10, magnetic moments are generated relative to the coordinate axes of the spacecraft - L _x , L _y , L _z ., which interacting with the vectors of magnetic induction of the Earth’s magnetic field B _x , B _y , B _z in accordance with formulas (2) create control moments M _x , M _y , M _z along the control axes of the spacecraft.

As magnetic actuators, for example, rod-type electromagnets with a core of magnetically soft material with windings of copper wire can be used (see A. P. Kovalenko. Magnetic control systems for spacecraft. M., "Engineering", 1975 ., p. 178). A description of the operation of reversible switches UPU _x , UPU _y , UPU _z - 5, 6, 7, for example, in the form of a bridge switch, is given in RF patent No. 2140128.

The duration of the damping process depends on the initial speed of separation of the spacecraft, the inertial characteristics of the spacecraft, the height of the orbit and the power of the magnetic drive (microcomputer control algorithms are data recorded in ROM).

Then, the reduction process and subsequent stabilization of the spacecraft in the orbital coordinate system OX _O Y _O Z _{O are} carried out. In this case, the axes of the orbital coordinate system OX _O Y _O Z _O and the axis of the associated spacecraft coordinate system OXsYsZs (the origin coincides with the center of mass of the spacecraft; the remaining axes are located along the spacecraft construction axes) must be aligned. The following angles are determined in the associated coordinate system of the spacecraft OXsYsZs:

- roll - rotation around the axis OXs;

- yaw - rotation around the axis OYs;

- pitch - rotation around the axis OZs

(with orbital orientation, the size of the body can be neglected).

The casting process starts automatically, or by command of the onboard computer.

Based on the navigation data periodically obtained from the on-board computer via IMI 16, the current position of the spacecraft center of mass in the geographical coordinate system is determined in relation to the current value of universal encoded time (UTS) and the geomagnetic induction vector is calculated according to the international geomagnetic field model IGRF (International geomagnetic reference field).

Since there is no network of supporting measurements of the secular course of the geomagnetic field in Russia, the world model of the normal field is used for interpretation purposes. This model is based on representing the normal field as a series of spherical functions, the coefficients of which are determined every 5 years on the basis of a global network of magnetic observatories (see RF patent No. 2447405).

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 - complement to latitude;

P _n ^m (cosθ _i ) is the associated Legendre function of the first kind;

g _n ^m , h _n ^m - spherical harmonic coefficients.

Based on the data obtained and using the transformation matrix from the geographic coordinate system to the orbital orientation system, the microcomputer 4 generates signals σ _x , σ _y , σ _z to form at the outputs of electromagnetic devices EM _x , EM _y , EM _z - 8, 9, 10 control magnetic moments relative to the coordinate axes of the spacecraft - L _x , L _y , L _z , providing a given control of the spacecraft for reorientation of the spacecraft to the orbital spacecraft, acceleration of the spacecraft in pitch to the orbital angular velocity.

The flywheel engine 14 (gyro stabilizer) is unwound to the rated speed, while electromagnetic devices EM _x , EM _y , EM _z - 8, 9, 10 constantly carry out correction of the spacecraft movement deviation arising due to the created kinetic moment of DM 14. The flywheel engine 14 is “frozen” into the spacecraft’s hull, i.e. does not make any movements relative to the body; at the same time participates in the movement of the latter. The characteristics of DM 14 are determined by its main parameter - the wearable kinetic moment (the maximum kinetic moment that DM 14 can accumulate). The flywheel engine 14 is used not only as a gyrostabilizer with a constant kinetic moment, but also for pitch control (within certain limits). Therefore, another selectable parameter is the maximum mechanical (control) moment, which DM 14 can create. The magnitude of the control torque DM 14 and the control of the value of its current rotation speed are carried out by the microcomputer 4. As a flywheel engine, for example, DM 1-20, made on the basis of a controlled torque non-contact direct current motor manufactured by JSC "VNIIEM Corporation", can be used. , Moscow, This flywheel engine has a pulse interface, an input for controlling the rotor speed and an integrated rotor speed sensor, by which the microcomputer 4 continuously detects the rotational speed torus DM 14. The interface device 15 provides the supply voltage to the DM 14 and pairing the signal from the output of the microcomputer 4, which determines the rotational speed of the rotor, thus setting the required values of the rotational speed of the flywheel DM 14.

Upon completion of the reorientation of the spacecraft and acceleration of the gyrostabilizer, the orientation system automatically (or by command of the microcomputer 4) switches to the normal mode - the main mode of operation to maintain the orientation and stabilization of the spacecraft in the orbital coordinate system:

- determination of the current position and speed of the spacecraft relative to the orbital coordinate system;

- bringing the spacecraft into the orbital coordinate system and maintaining this orientation with a given accuracy.

In this mode, the flywheel engine 14 works primarily as a gyrostabilizer that allows you to resist the action of the disturbing moment and “holds” the orientation in those parts of the orbit in which the induction vectors of the geomagnetic field and the disturbing moment are almost parallel.

The speed of rotation of the flywheel DM 14 should be a constant stabilized value. The maintenance of this mode is ensured by the microcomputer 4, by monitoring the frequency using the built-in rotor speed sensor DM 14.

During the motion of a spacecraft oriented in a certain way to the Earth, to maintain this orientation, it must constantly rotate in pitch. To do this, changing the frequency of rotation of the flywheel relative to the stabilized speed, creates a kinetic moment K in pitch and provides a given angular motion of the spacecraft's hull. The return to the stabilized speed of DM 14 and the reset of the kinetic moment is carried out by the second circuit of the system of executive bodies, capable of creating external control moments, namely, electromagnetic devices EM _x , EM _y , EM _z - 8, 9, 10 (the formation of the controlling mechanical moment is discussed above) .

The full kinetic moment G can be represented as the sum of two kinetic moments: the spacecraft K body and the gyrostabilizer Н

The second circuit of the system of executive bodies operates in the following mode. As soon as the magnitude of the vector of kinetic momentum H developed by the gyrostabilizer increases due to the increase in kinetic momentum G, the external control moment created by the second circuit of the system of executive organs will reduce G, as a result of which, in accordance with formula (4), the required value of N. will decrease. the actual value of H will be achieved by returning the gyrostabilizer operation mode to the previous one, at which the value of H has not yet managed to grow. Due to the continuous inclusion of the second circuit and additional adjustment of the rotor speed, the microcomputer 4 keeps the speed of the flywheel DM 14 stable within acceptable limits.

Thus, high reliability of the spacecraft orientation system with a given accuracy is achieved.

This ensures high reliability of telemetry, and microprocessor control of the operating modes of the device, the use of one reliable flywheel engine, a reliable active magnetic orientation system can provide technical and operational requirements necessary for small spacecraft (especially for multi-tiered grouping):

- minimum weight, dimensions, power consumption and cost;

- high reliability of the spacecraft (the active life of the MCA is at least 5 years);

- maximum simplicity;

- high technology;

- minimum time for ground mining, testing and preparation for launch.

Claims (2)

1. The method of orientation of the spacecraft, which consists in measuring the magnetic induction of the Earth's magnetic field along the three coordinate axes X, Y, Z, forming control signals for electromagnetic devices, forming magnetic moments at the outputs of electromagnetic devices on each axis of the spacecraft, characterized in that the kinetic moments are determined of the inertial mechanical system “KA-flywheel housing” in roll, yaw, pitch and the flywheel engine in pitch, according to the received data of the kinetic moments of the inertial mechanical system “KA-flywheel housing” ik ”carry out preliminary reassurance and control of the spacecraft by roll, yaw, pitch with the help of electromagnetic devices, maintain a constant rotational speed of the rotor of the flywheel engine to use it as a gyro stabilizer, further regulate the speed of the flywheel engine by controlling the rotor speed using microcomputers to maintain the required angular velocity of the spacecraft in pitch, determine the kinetic moment of the flywheel engine and reset it using electromagnetic devices. 2. The orientation device of the spacecraft, which includes three sensors of the magnetometer, the outputs of which are connected to the input device of the microcomputer, the output devices of which are connected to the inputs of three amplification-converting devices, the outputs of which are connected to the inputs of three electromagnetic devices, characterized in that three component sensors are introduced kinetic moment, the outputs of which are connected to the input devices of the microcomputer, the engine-flywheel, the information interface module and the interface device, the input of which is connected to the device water of the microcomputer, and the output - with the control input of the flywheel engine, the output of the rotor speed sensor of which is connected via the information transmission line to the information interface module, which is connected to the central computer and microcomputer via the information transmission line.

Images