RU2812876C1 - Astronomical orientation system for orbital spacecraft with feedback - Google Patents

Abstract

FIELD: space technology; spacecraft orientation systems.

SUBSTANCE: celestial orientation system of an orbiting spacecraft (SC) contains a star sensor (SS), a module for calculating navigation and ballistic information (MCNBI), a module for calculating a stabilization matrix (MCSM), a module for calculating stabilization angles (MCSA) and a block of gyroscopic angular velocity meters (BGAVM). The outputs of SS and MCNBI are connected to the first and second inputs of MCSM, the output of which is connected to the input of MCSA. The system includes: a module for calculating the parameters of the spacecraft's programmed motion (MCPSPM), a module for generating a stabilization matrix (MGSM), a module for compensating for the mutual influence of channels (MCMIC) and a module for generating optimizing program functions (MGOPF). Three integrators, three amplification-matching units (AMU) and nine adders have also been introduced. Each of the MCSA outputs forms separate circuits of series-connected first adders, AMU, second adders, third adders and integrators, the outputs of which are connected to the second inputs of the first adders and the corresponding inputs of the MGSM. The second inputs of the second group of adders are connected to the corresponding outputs of the BGAVM, and their third inputs are connected to the corresponding outputs of the MCPSPM, which are simultaneously connected to the corresponding inputs of the MCMIC. The first, second and third outputs of the MCMIC are connected to the corresponding second inputs of the third group of adders. The fourth output of the MCPSPM is connected to the third input of the MCSM, and its first and second inputs are connected to the second output of the MCNBI and the output of the MGOPF. The outputs and inputs of the integrators in each of the channels are the angles of stabilization of the spacecraft in terms of angles and angular velocities, respectively.

EFFECT: invention makes it possible to improve the quality of spacecraft control in orbital flight.

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Description

The invention relates to the field of space technology and can be used to orient a spacecraft (SV) relative to the orbital (OSC), program (PSC) and inertial (ISK) coordinate systems using a star sensor (DS).

There are numerous works, for example [1-15], where the issues of celestial orientation of spacecraft in near-Earth space are considered.

In particular, in the book [4] by V.N. Branets, I.P. Shmyglevsky “Application of quaternions in problems of orientation of a rigid body.” Moscow, Nauka 1973 (see pp. 205-226), some block diagrams of orbital astro-orientation devices are given, while the authors consider only general theoretical aspects of spacecraft orientation.

For example, the device [14] is known, discussed in the article “System for orientation and stabilization of a spacecraft based on information from astro sensors” // Electronic journal “Proceedings of MAI”. Issue No. 38, which presents the results of flight tests, but does not insufficiently disclose the essential features of the attitude control system.

In the popular book Spacecraft Systems Development//Edited by P. Fortescue, G. Swinerd, J. Stark; Per. from English-M.: Alpina publisher, 2015, issues of spacecraft reorientation are considered, also without specifics of circuit solutions.

Similar shortcomings are contained in sources [1-2, 5-13, 15] and many others.

The closest device that can be taken as a prototype is the device described in article [3] by authors K.A. Boyarchuk et al. Orientation and stabilization system of the Condor-E spacecraft. Proceedings of section 22 named after academician V.N. Chelomeya XXXVIII academic readings on astronautics. The indicated attitude control system (OS) uses the principle of astrocorrection with feedback, however, the AO is constructed in such a way that it is limited only by programmatic directional rotation without the possibility of programmatic rotations of the spacecraft along other axes - roll and pitch. At the same time, the programmed motion itself is not stabilized during the entire time of the programmed rotation, which leads to a low quality of the transition process and, as a consequence, a decrease in the accuracy of the spacecraft orientation.

The technical result of the invention is to improve the quality of spacecraft control in orbital flight - the implementation of three-dimensional program orientation (including pitch and roll) and the optimization of transient processes of the spacecraft angular motion parameters.

For this, in contrast to the known technical solution, which contains a star sensor (DS), a module for calculating navigation and ballistic information (MRBI), a module for calculating a stabilization matrix (MRMS) and a module for calculating stabilization angles (MSA), and the outputs of the DS and MRBI are connected to the first and second inputs of the MRMS, the output of which is connected to the input of the MRU, as well as a block of gyroscopic angular velocity meters (GYUS), new connections and new software modules have been introduced into it - a module for calculating the parameters of the programmed motion of the spacecraft (MRPPD), a module for generating a stabilization matrix (MFMS), a module for compensating for the mutual influence of channels (ICVC) and a module for generating optimizing software functions (MFOPF), as well as three integrators, three amplification-matching units (USB) and nine adders, with each of the MRU outputs forming separate circuits from sequentially connected first adders, USB, second adders, third adders and integrators, the outputs of which are connected to the second inputs of the first adders and the corresponding inputs of the MFMS, the second inputs of the second group of adders are connected to the corresponding outputs of the BIUS, and their third inputs are connected to the corresponding outputs of the MRPPD, which simultaneously connected to the corresponding inputs of the ICVC, the first, second and third outputs of the ICVC are connected to the corresponding second inputs of the third group of adders, in addition, the fourth output of the MRPPD is connected to the third input of the MRMS, and its first and second inputs are connected to the second output of the MRNBI and the output of the MFOPF, and the outputs and the inputs of the integrators in each of the channels are the angles of stabilization of the spacecraft in terms of angles and angular velocities, respectively.

In fig. 1 shows an example of the invention, where it is indicated:

- orientation matrix of the spacecraft relative to the ISK, generated by the remote sensing;

- orientation matrix of the USC relative to the ISK, formed by MRNBI;

- matrix of programmed rotations of the spacecraft by angles

C* - stabilization matrix according to remote sensing data, MRNBI and MRPPD;

- calculated angles of mismatch between the UCS and SSC;

- angles and angular velocities of spacecraft stabilization supplied to the executive bodies, respectively, in pitch, heading and roll;

C - matrix formed in feedback circuits according to the spacecraft stabilization angles - Δϑ, Δψ, Δγ;

1, 2, 3 - amplifying-matching blocks in the pitch channel-1 (K ₃ ), in the heading channel - 2 (K ₂ ), in the roll channel - 3 (K ₁ );

γ _y , ψ _y , ϑ _y - program rotation settings transmitted to the spacecraft via KRL, relative to the USC, respectively, in roll, heading and pitch;

- optimizing software functions for the angles and angular velocities of the angular motion of the spacecraft, respectively, for roll, heading and pitch;

- angular velocities of the UCS relative to the ISK;

- angular velocities of the USC relative to the ISK;

p, q, r - BIUS readings relative to the SCS axes (X, Y, Z), respectively.

Everywhere - roll, - well, - pitch.

4, 5, 6 - first, second and third integrators in the pitch, heading and roll channels, respectively;

1.1-1.3 - first group of adders;

2.1-2.3 - second group of adders;

3.1-3.3 - the third group of adders.

The celestial orientation system works as follows.

The MRMS module calculates the current value of the stabilization matrix (mismatch between SSC and PSC) - C*. It is calculated:

- according to remote sensing readings -

- according to ballistic calculation data on the position of the USC relative to the ISK - A (Ω, i, u);

- according to the current values of the matrix of the required program rotation of the spacecraft relative to the OSK -

The MRU, having received the matrix C*, calculates the mismatch angles between the UCS and SSC:

The calculated angles ϑ _cn , ψ _cn , γ _cn, each in its own channel, are compared with the spacecraft stabilization angles - Δϑ, Δψ, Δγ, which are obtained at the output of integrators 4, 5, 6:

These signals are amplified in the corresponding CSS - 1, 2, 3 and added with the CICS signals (p, q, r) on the corresponding adders of the second group - 2.1, 2.2, 2.3.

Next, the signals arrive at the corresponding first inputs of the adders of the third group 3.1, 3.2, 3.3, where the corrections coming from the ICVC for “interfering” cross-connections of program rates are subtracted from them:

and after integration on the corresponding integrators - 4, 5, 6 are supplied to the corresponding second inputs of the first group of adders (1.1, 1.2., 1.3) with opposite signs, thereby closing the feedback at the stabilization angles.

To generate corrections, the ICVC module receives data on program angular velocities from the MRPPD and matrix C, composed of the actual spacecraft stabilization angles - Δϑ, Δψ, Δγ, coming from the outputs of integrators 4, 5, 6:

the result is the formation of ICWC signals:

The MRPPD module generates programmed speeds of the spacecraft relative to the ISK - receiving data from the MFOPF module in the form of program angles and program angular velocities - and from the MRNBI module in the form of the angular velocities of the USC relative to the ISK -

Where - matrices of specified flat program rotations of the spacecraft along the heading, pitch and roll angles, respectively

Where - precession speed of the line of USC nodes;

- the angle of inclination and the rate of change in the angle of inclination of the USC plane to the plane of the Earth's equator;

- the argument of latitude and the rate of change of the argument of latitude (orbital angular velocity - ω _O ).

The MFOPF module converts the settings coming from the flight control center (MCC) via the command radio link (CRL), into the program angles and program angular velocities of the spacecraft -

This is necessary due to the fact that during the program rotations of the spacecraft, the specified settings and angles of the program rotations of the spacecraft are not equal and they will coincide only at the end of program turns after the attenuation of transient processes. At the same time, the MFOPF ensures stabilization of the angular velocities of the spacecraft relative to the programmed trajectories of its angular motion (Fig. 2, 3), which makes it possible to perform a programmed rotation of the spacecraft with the required quality of the transient process.

In fig. Figure 4 shows the process of programmed turns of the spacecraft without stabilization of angular velocities relative to the programmed trajectories - the software rotation process turns out to be of poor quality, with a very large overshoot of ~ 10°.

In accordance with the structural and functional diagram (Fig. 1), the equations of motion of the celestial orientation system have the form:

- matrices of specified flat program rotations of the spacecraft along the heading, pitch and roll angles, respectively

Functionality of MFOPF:

The implementation of the functionality is shown in Fig. 2, 3.

Functional forms control in the form of a general trapezoid; as a result, transient processes of programmed rotations of the spacecraft for each of the orientation angles are performed in a minimum amount of time and have an almost ideal appearance.

Note that the celestial orientation system with feedback does not require the use of quaternions or a final rotation vector, but at the same time allows you to perform programmed rotations of the spacecraft at unlimited angles simultaneously - along all axes (heading, pitch, roll) - fig. 5. We also note that celestial orientation with feedback makes it possible to obtain extremely small orientation errors:

Where - astro sensor errors, - drift of BIUS gyroscopes. For example, for And spacecraft orientation errors will be:

Thus, the system of celestial orientation of an orbital spacecraft with feedback makes it possible to significantly improve the quality of control of the angular motion of the spacecraft in orbital flight through the implementation of three-dimensional program orientation and optimization of transient processes of the parameters of the program angular motion of the spacecraft.

Literature.

1. Alekseev K.B., Bebenin G.G. Control of spacecraft//M.: Mechanical Engineering. 1974

2. Anuchin O.N., Komarova I.E., Perfilyev L.F. On-board navigation systems and orientation of artificial Earth satellites // St. Petersburg: State Scientific Center of the Russian Federation Central Research Institute "Electropribor". 2004

3. Boyarchuk K.A. and others. Orientation and stabilization system of the Condor-E spacecraft // Proceedings of section 22 named after academician V.N. Chelomeya XXXVIII-x Academic readings on cosmonautics. 2015. T22. S408-434.

4. Branets V.N., Shmyglevsky I.P. Application of quaternions in problems of orientation of a rigid body // M.:.: Nauka 1973.

5. Vasiliev V.N. Spacecraft orientation systems // M.: FSUE “NPP “VNIIEM”. 2009.

6. Kavinov I.F. Inertial navigation in near-Earth space// M.: Mashinostroenie. 1988

7. Kargu L.I. Systems for angular stabilization of spacecraft // M.: Mashinostroenie. 1980

8. Quasius G., McCanless F. Design of celestial navigation systems // M. Mir. 1970.

9. Kochetkov V.I.. Systems of astronomical orientation of spacecraft // M.: Mechanical Engineering, 1980.

10. Raushenbakh B.V., Tokar E.N. Spacecraft orientation control//M.: Nauka. 1974

11. Development of spacecraft systems. Edited by P. Fortescue, J. Stark, G. Swinerd. M.: Alpina publisher. 2017

12. Solovyov V.A., Lysenko L.N., Lyubinsky V.E. Space flight control // Moscow Higher Technical School named after. N.E. Bauman. 2009

13. Solodov A.V. Engineering reference book on space technology//M.: publishing house of the USSR Ministry of Defense. 1969

14. System of orientation and stabilization of a spacecraft based on information from astro sensors / Electronic journal “Proceedings of MAI”. Issue No. 38. 2017

15. Modern problems of orientation and navigation of spacecraft // Space Research Institute of the Russian Academy of Sciences. Collection of works 2012

Claims (1)

The celestial orientation system of an orbiting spacecraft (SC), containing a star sensor (DS), a module for calculating navigation and ballistic information (MRBI), a module for calculating a stabilization matrix (MRMS) and a module for calculating stabilization angles (MSA), and the outputs of the DS and MRBI are connected to the first and the second input of the MRMS, the output of which is connected to the input of the MRU, as well as a block of gyroscopic angular velocity meters (GYUS), characterized in that it contains: a module for calculating the parameters of the programmed motion of the spacecraft (MRPPD), a stabilization matrix generation module (MFMS), a module channel interference compensation (ICVC) and a module for generating optimizing software functions (MFOPF), as well as three integrators, three amplification-matching units (USB) and nine adders, with each of the MRU outputs forming separate circuits of series-connected first adders, ASB, second adders, third adders and integrators, the outputs of which are connected to the second inputs of the first adders and the corresponding inputs of the MFMS, the second inputs of the second group of adders are connected to the corresponding outputs of the BIUS, and their third inputs are connected to the corresponding outputs of the MRPPD, which are simultaneously connected to the corresponding inputs of the ICVC, the first, second and third outputs of the ICVC are connected to the corresponding second inputs of the third group of adders, in addition, the fourth output of the MRPPD is connected to the third input of the MRMS, and its first and second inputs are connected to the second output of the MRNBI and the output of the MFOPF, and the outputs and inputs of the integrators in each of the channels are the angles of stabilization of the spacecraft in terms of angles and angular velocities, respectively.

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