RU2517018C2 - Method for automatic compensation of strapdown attitude reference system errors in spacecraft orientation control system and apparatus realising said method - Google Patents

Abstract

FIELD: physics, navigation.

SUBSTANCE: group of inventions relates to spacecraft strapdown attitude reference systems (BSO) with gyro-inertial and astro-navigation elements The disclosed method involves compensating BSO errors caused by systematic errors of angular velocity sensors. The method is based on comparing readings of angular velocity sensors and astroorientation sensors. An estimate of the angular velocity sensor error is generated by isodromic transformation of the result of comparing signals of the angular position of the spacecraft, measured by an astroorientation sensor unit and calculated in a BSO computer based on the angular velocity sensor readings. The angular velocity sensor error estimate signal at the output of the isodromic link is generated while connecting to the astroorientation sensor unit control. That signal is stored over the duration of disconnection from the astroorientation sensor unit control. The obtained signal is simultaneously subtracted from the angular velocity sensor readings continuously. Said angular velocity sensor readings reflect the absolute angular velocity of the spacecraft and the error caused by systematic errors. The apparatus which realises the disclosed method comprises a BSO, an astroorientation sensor unit and actuating devices. The gyro-inertial measurement unit of the BSO consists of single-component angular velocity sensors. A computer integrates kinematic equations based on information from the angular velocity sensors. Computer modules include isodromic links, adders and comparator elements.

EFFECT: high accuracy of spacecraft orientation in continuous operating mode of the spacecraft owing to constant automatic compensation of BSO errors.

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Description

FIELD OF THE INVENTION

The invention relates to space technology and can be used in a spacecraft (SC) orientation control system based on a strapdown orientation system.

State of the art

Analogs of orientation control systems, as a rule, contain angular position sensors, angular velocity sensors, computing devices or on-board computers and executive bodies (Raushenbakh B.V., Tokar E.N. Spacecraft orientation control. - Moscow: Nauka, 1974 ; Alekseev KB, Bebenin GG Control of spacecraft. - Engineering, 1974; Vasiliev VN. Orientation systems for spacecraft. - M.: FSUE NPP VNIIEM, 2009).

Currently, in connection with the improvement of on-board computers and the use of high-precision angular velocity sensors in spacecraft orientation control systems, a control method is used in which the orientation parameters are calculated and control signals are generated by the executive bodies on the basis of strapdown orientation systems. In this case, the signals of the angular mismatches of the strapdown orientation systems for determining the orientation of the spacecraft can be obtained by calculations in the computing devices for measuring the angular velocity using angular velocity sensors by continuously integrating the kinematic equations of the angular motion of the spacecraft (see Branets V.N., Shmyglevsky I.P. Introduction to the theory of strapdown inertial navigation systems. - M .: Nauka, 1992; Pelpor D.S. Gyroscopic systems. - M .: Higher school, 1988).

The disadvantage is the need for external error correction of the strapdown orientation system, since integration of the kinematic equations of the angular motion of the spacecraft inevitably accumulates errors from the errors of gyro-inertial sensors of angular velocities and calculation errors. Compensation of errors in the strapdown orientation system is usually carried out using the readings of accurate astro-orientation sensors.

A known method of orienting a spacecraft and a device that implements this method. according to the patent for the invention of the Russian Federation No. 2375269. The method includes measuring the parameters of the angular motion of the spacecraft, forming and issuing control signals to the executive bodies. At the same time, they provide a cyclic calculation of the time instants of subsequent cycles of the formation and generation of spacecraft control signals. For the reference period of the current time, the equipment is on and the equipment for normal operation modes of the orientation system is turned off. If the current time coincides with the estimated time of the next cycle, turn on the standard equipment and turn off the equipment of standby operation modes. Measure the parameters of the angular motion of the spacecraft, form and issue the necessary control signals to the executive bodies. The spacecraft orientation device can be made in the form of a strapdown inertial control system. It contains a central control unit, angular velocity meters, external information sensors and executive bodies. In addition, control units for angular velocity meters, control units for external information sensors, and a program-time device have been added. These blocks and the device are connected to each other and other elements of the control system in accordance with the implemented method.

The disadvantage is the cyclical operation in normal mode, that is, the desired orientation accuracy is not provided during continuous operation.

The closest technical solution is a method for calibrating gyro-inertial meters of the strapdown spacecraft orientation system according to RF patent No. 2092402, chosen by the authors for the prototype. The method consists in calibrating spacecraft turns around the roll and pitch axes with the sighting of the given astro-orientations before the start of each turn and the subsequent calculation of the error of the gyro-inertial gauge unit composed of one-component angular velocity sensors, as well as the compensation of the influence of the constant departures of the gyro-inertial gauges (angular velocity sensors) as a component of the error in their output. In this case, the calibration turns are carried out by three consecutive plane rotations of the spacecraft around the roll, yaw and pitch axes of the coordinate system associated with the object of orientation by a given angle, with the sight of two specified astronomical points using astro-orientation sensors before and after each of the spacecraft’s calibration turns, after which according to the measuring information obtained as a result of astrovization and information obtained using the strapdown orientation system, the error values are estimated sensors of angular velocity.

The system that implements this method contains a block of astroorientation sensors, a storage device designed to store a catalog of astroorientations, a computing device for determining the astroorientation of a spacecraft, a block of gyroinertial meters, composed of three one-component angular velocity sensors, a strapdown inertial system that integrates kinematic equations according to information about absolute the angular velocity of the orientation object, the executive bodies performing the calibration spreads You are the spacecraft around the roll, yaw and pitch axes of the coordinate system associated with the spacecraft.

The disadvantage is the need for sufficiently long calibration rotations of the spacecraft. Given that the parameters of gyroscopes during operation can change, for example, due to degradation, the calibration turns should be repeated regularly, while the required accuracy of orientation in the continuous operation of the spacecraft is not ensured. In addition, the loss of time for calibration rotations can create significant inconvenience to the target use of the spacecraft and, in addition, in these modes lead to increased power consumption by the spacecraft orientation control system.

Disclosure of invention

The aim of the invention is to provide high accuracy control the orientation of the spacecraft in a continuous mode of operation of the spacecraft due to the constant automatic compensation of errors.

This goal is achieved by the fact that they provide continuous automatic compensation for errors of the strapdown orientation system caused by low-frequency components of the systematic errors of the angular velocity sensors, which are the main sources of errors.

The application of this method provides the necessary accuracy of the spacecraft orientation control systems, while, unlike the prototype, calibration spacecraft rotations are not required, which means the spacecraft’s useful life is increased and its energy consumption is reduced.

The essence of the invention lies in the fact that the proposed method for automatically compensating for errors of the strapdown orientation system included in the spacecraft orientation control system (caused by systematic errors of the gyro-inertial measuring unit composed of single-component angular velocity sensors) is based on comparing the readings of the angular velocity sensors and the astro-orientation sensor block, established on a common base and oriented in a certain way relative to the axes of the associated coordinate system INAT spacecraft. In this case, an error estimation signal of the angular velocity sensors is used, which is generated by isodromic conversion of the signal obtained by comparing signals characterizing the angular position of the spacecraft measured by the astroorientation sensor unit with the angular position calculated in the computing device of the strapdown orientation system according to the readings of the angular velocity sensors. The signal for estimating the error of the angular velocity sensors at the output of the isodromic link is generated when the astroorientation sensor unit is connected to the control unit, it is stored for the time that the astroorientation sensor unit is disconnected from the control unit, and at the same time, the obtained signal for evaluating the error in the error of the angular velocity sensors from the readings of the angular velocity sensors is continuously which reflect the absolute angular velocity of the spacecraft and the error caused by the systematic errors of the angular velocity sensors.

Thus, they automatically and continuously compensate for the errors of the strapdown orientation system caused by the systematic errors of the angular velocity sensors, both during connection to the control and during disconnection of the astro-orientation sensor block from the control.

The goal is realized by the device of the spacecraft orientation control system by including additional computing modules with appropriate connections in its composition to ensure continuous automatic compensation of errors caused by low-frequency components of the systematic errors of angular velocity sensors.

The spacecraft orientation control system contains a strapdown orientation system, a block of astro-orientation sensors, and executive bodies. In this case, the strapdown orientation system includes a block of gyroinertial meters composed of one-component angular velocity sensors, and a computing device with integrating links that integrates kinematic equations using information about the absolute angular velocity of the spacecraft. The astroorientation sensor unit contains astro-vision devices, a storage device designed to store the astro-orientation catalog, and a computing device for determining the astro-orientation of the spacecraft. Additionally, computational modules (isodromic links, consisting of an integrating link and a linear amplifier connected in parallel, adders and comparison elements installed on each axis of the associated coordinate system of the spacecraft) introduced into the strap-down orientation system provide constant tracking of deviations caused by low-frequency components of the systematic errors of the angular velocity sensors , and compensate for these deviations. This ensures high accuracy of the spacecraft orientation control system. At the same time, the calibration rotations of the spacecraft are not required, which means that, compared with the prototype, the useful life of the spacecraft is increased and energy consumption is reduced.

The drawing shows a simplified block diagram of a satellite orientation control system, which is a device that implements the inventive method.

Numbers denote the following elements:

1 - strapdown orientation system (BSO);

2 - block gyroinertial meters (BHII);

3 - a block of astro-orientation sensors;

4 - element of comparison;

5 - isodromic link

6 - adder;

7 - computing device (WU);

8 - on-board computer (BC);

9 - block coordination and control commands (BSU);

10 - executive bodies (IO);

11 - regulatory object (spacecraft - SC);

12 - central on-board computer (PPM)

The implementation of the invention

The spacecraft orientation control system shown in the drawing comprises a strap-on orientation system 1; astro-orientation sensor block 3 (containing astro-sighting devices not shown in the figure, a storage device designed to store the astro-orientation catalog, and a computing device for determining the astro-orientation of the spacecraft), as well as executive bodies 10.

The strapdown orientation system 1 contains a block of gyroinertial meters 2, composed of three one-component angular velocity sensors, which are installed on a common base with astrovisir devices of the astroorientation sensor block 3 and are oriented in a certain way relative to the axes of the associated spacecraft coordinate system. Also contains series-connected comparison element 4, isodromic link 5, adder 6; computing device 7 with integrating links (here computing device 7 consists of two units: on-board computer 8 and the coordination unit and control commands 9). The drawing conventionally shows the regulatory object - KA 11 and the central on-board computer 12, which is not included in the spacecraft orientation control system, but control commands and information from it are used in the spacecraft orientation control system in accordance with its operation algorithm.

In the embodiment of the spacecraft orientation control system shown in the drawing, the output of the matching unit and control commands 9 of the computing device 7 is connected to the actuators 10 and to the comparison element 4, the second input of the comparison element 4 is connected to the output of the astro-orientation sensor block 3. The output of the comparison element 4 is connected to the input of the isodromic link 5, the output of the isodromic link 5 is connected to one input of the adder 6, the second input of which is connected to the output of the gyro-inertial measuring unit 2, the output of the adder 6 is connected to matching block and the move commands 9 which is interconnected with the board computer 8. The central on-board computer 12 is connected with the onboard computer of the computing device 7 and a unit astroorientation 3 sensors.

In addition, the structural diagram of the spacecraft orientation control system indicates: w is the angular velocity vector at the output of the block of gyro-inertial amplifiers, w _a is the absolute angular velocity vector of the spacecraft, Δ is the systematic error vector of the TLS, φ _a is the angle vector measured by the astro-orientation sensor unit, φ is the angle vector calculated in the strapdown orientation system, ε is the mismatch signal vector.

In the presented embodiment of the spacecraft orientation control system, as a unit of gyroinertial meters 2, a strapdown gyro-orientation device (BGO device) manufactured by NIIEM can be used.

As a block of astro-orientation sensors 3, a star sensor (BOKZ-M device) manufactured by IKI RAS can be used.

As an on-board computer 8, the BTsVM-K device manufactured by the Submicron Research Institute can be used.

As a coordination unit and control commands 9, a BSK device manufactured by the Submicron research institute can be used.

As the executive bodies 10 can be applied engines - flywheels (DM5-50 devices) produced by VNIIEM.

The strapdown orientation system 1 can be implemented in the form of blocks of an onboard computing device, which will include, with their connections, a comparison element 4, an isodromic link 5, an adder 6, an onboard computer 8, and a coordination unit and control commands 9.

During the normal operation of the spacecraft orientation control system from the central on-board computer 12, commands are sent to the astro-orientation sensor unit 3 and to the on-board computer 8 to set the orientation modes, navigation data and other information to provide a predetermined mode for the operation of the orientation control system. In the astro correction mode (with the astroorientation sensor unit connected), the signal φ _a , which characterizes the angular position of the spacecraft measured by the astroorientation sensor unit in vector form, is received from the astroorientation sensor unit 3 to the input of the comparison element 4. At the same time, a signal φ is received at the second input of the comparison element 4, which characterizes the angular position of the spacecraft in a vector form, calculated in the matching unit and control commands 9 of computing device 7 by integrating the kinematic equations of the angular motion of the spacecraft. In addition, the coordination unit and control commands 9 carries out electrical and informational interaction between sensors and executive bodies in the control channels, and the implementation of all computational algorithms related to the spacecraft orientation control system is performed on-board computer 8 and in the coordination unit and control commands 9, between which there is a continuous exchange of information. At the same time, the angular velocity sensors of the unit of gyroinertial meters 2 measure the projections of the angular velocity of the spacecraft on the axis of the associated coordinate system, and from the output of this block, the signal w = w _a + Δ, which characterizes the angular velocity of the spacecraft in vector form, where w is the angular vector the velocity at the output of the BGII 2, w _a is the vector of the true angular velocity of the spacecraft, Δ is the vector of the systematic error of the angular velocity sensors. From the output of the adder 6, the signal is fed to the integrating links of the matching unit and control commands 9 of the computing device 7, where the signal φ is generated. The signal φ, as noted above, is supplied to the comparison element 4 and at the same time to the executive bodies 10. The executive bodies 10 act on the regulatory object (KA) 11. The KA processes the signal φ, approaching the mismatch ε = φ _а -φ at the output comparison element 4. The signal from the comparison element 4 is fed to the isodromic link 5. The isodromic link 5 accumulates at the output a low-frequency component of the error estimation of the angular velocity sensors until the signal at the input of the isodromic link differs from zero, stores the accumulated nnuyu low-frequency component of the angular velocity sensor evaluation and stores it on the operation strapdown system orientation astrocorrection mode and time astrocorrection interrupt. The error estimation signal obtained after isodromic conversion of the angular velocity sensors then goes to the adder 6, where it is subtracted from the signal of the gyroinertial measuring unit 2, which contains the absolute angular velocity of the spacecraft and the error caused by the systematic errors of the angular velocity sensors. Thereby, the low-frequency components of the systematic errors of the angular velocity sensors are automatically compensated both during the connection of the astro-orientation sensor block to the control and during disconnection of the astro-orientation sensor block from the control.

Thus, the proposed device and the method for its implementation automatically and continuously compensate for the low-frequency components of the systematic errors of the angular velocity sensors, which leads to an increase in the accuracy of the orientation of the spacecraft, while, unlike the prototype, calibration rotations of the spacecraft are not required, and thereby increase the time beneficial use of spacecraft and reduced power consumption.

The proposed spacecraft orientation control system and the automatic BSO error compensation method as part of the spacecraft orientation control system have industrial applicability and are currently being implemented in the spacecraft orientation control system under development, the prototypes of which are being tested.

According to the test results of the prototypes, the spacecraft orientation control system provides an error in determining the orientation (3σ) in the astrocorrection mode on all axes of not more than 1 angle. min, in the "gyro-memory" mode after performing astro correction for 100 min, not more than 1 angle. min

Claims (2)

1. A method for automatically compensating for errors of a strapdown orientation system in a spacecraft’s orientation control system, which consists in compensating for the influence of constant departures of gyroinertial meters as an error component in their output signal, based on a comparison of the readings of angular velocity sensors and astroorientation sensors, characterized in that they use an estimation signal errors of the angular velocity sensors, which are formed by isodromic conversion of the signal obtained as a result of introducing signals characterizing the angular position of the spacecraft, measured by the block of astroorientation sensors, with the angular position calculated in the computing device of the strapdown orientation system according to the readings of the angular velocity sensors, while the signal for estimating the error of the angular velocity sensors at the output of the isodromic link is formed during connection to the control unit sensors of astro-orientation, remember it for the time of disconnection from the control of the block of sensors of astro-orientation and at the same time carry out the continuous subtraction of the received signal for estimating the error of the angular velocity sensors from the readings of the angular velocity sensors, which reflect the absolute angular velocity of the spacecraft and the error caused by the systematic errors of the angular velocity sensors. 2. The control system for the orientation of the spacecraft, containing a strap-on orientation system, a block of astroorientation sensors, executive bodies, while a strap-down orientation system contains a block of gyro-inertial meters made up of one-component angular velocity sensors, and a computing device with integrating links, a block of astro-orientation sensors contains astrovisir devices , a storage device for storing a catalog of astro-landmarks, and a computing device your definition of astro-orientation of the spacecraft, characterized in that it contains additional computing modules with corresponding connections, namely, an isodromic link, an adder and a comparison element are introduced into the strapdown orientation system on each axis of the connected coordinate system of the spacecraft, while the output of the computing device of the strapdown orientation system is connected to the executive bodies and to the comparison element, the second input of the comparison element is connected to the output of the sensor block astroorientation ii, the output of the comparison element is connected to the input of the isodromic link, the output of the isodromic link is connected to the input of the adder, the output of the angular velocity sensor of the gyro-inertial measuring unit is connected to the second input of the adder, and the output of the adder is connected to the input of the computing device of the strapdown orientation system.

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