Method of three-axis orientation of spacecraft in orbital coordinate system

FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

 

The invention relates to space technology and can be used in the design of traffic management systems (DMS) SPACECRAFT).

It is known that the direct use of the information on the position of the Sun produced by the device orientation by the Sun (PIC), the traffic management system, maintaining constant the position of the Sun in the field of view of the instrument SETTLEMENT, provides orientation to one of the associated SPACECRAFT axes or associated with the AC vector in the direction of the Sun.

From the patent literature known way triaxial orientation of the SPACECRAFT in the orbital coordinate system according to the device orientation to the Sun (see, e.g., auth. St. No. 1655842, CL 64 G 1/00 from 02.12.88 year).

However, not clearly defined the position of the SPACECRAFT during the turn around pointed at the Sun associated with the KA axis or centered on the Sun associated with SPACECRAFT vector, i.e. not clearly provided tricosta orientation and adjustment related to SPACECRAFT axes in space.

The objective of the invention is to provide a method triaxial orientation of the spacecraft in the orbital coordinate system (USC) with achievement of the technical result in a reduction of weight, volume, and simplify the design of SPACECRAFT by reducing the number of devices on Board the SPACECRAFT during the implementation of the three-axis is rantatie, as well as expanding Arsenal of technical tools in this field of technology.

This task is solved in that way triaxial orientation of the spacecraft in the orbital coordinate system according to the device orientation to the Sun, in accordance with the invention, the onboard computer enter orbit parameters, calculate the Sun position in the field of view of the instrument orientation for each point of the orbit to the orientation associated with the spacecraft axis in the orbital coordinate system, define a space apparatus search engine angular velocity to ensure capture of the Sun field of view of the instrument orientation, then reduce down to zero angular velocity, ensuring the presence of the Sun in the field of view of the instrument orientation, then produced a roll of the spacecraft so to the Sun in the hearth of the review of the device orientation has moved to the desired initial calculated from the positions of the Sun in the field of view of the device orientation and then continuously deploy spacecraft from the starting point in the following the calculated position of the Sun in the field of view of the instrument orientation for each point of the orbit at the desired orientation associated with spacecraft axis in the orbital coordinate system.

Hereinafter the invention is explained using figures.

In Fig., explaining the position of the Sun in the field of view of the instrument PIC shown (Fig.2 and 3) the position of the Sun during the motion of SPACECRAFT in orbit in oriented relative to the RSC position associated with the SPACECRAFT axes, which can be calculated for each point of the orbit.

Figure 1 shows the following notation:

1 - the field of view of the device POS;

X, Y, Z - linked axis KA;

S - the position of the Sun in the field of view after grabbing him by the POS device (arbitrary);

α, β Sun angles produced by the PIC;

S1array calculated by polozhenie of the Sun in the field of view of the instrument PIC when moving SPACECRAFT in orbit from point 1 to point 2 (see figure 2 and 3) when the orientation associated with the SPACECRAFT axes at USC;

Point 1, point 2 - point start and end of the calculated positions of the Sun during the motion of SPACECRAFT in orbit from point 1 to point 2.

In figure 2, illustrating the position of the orbit and associated SPACECRAFT axes relative to the Sun at the pole of the World, given the following notation:

X, Y, Z - linked axis KA;

Point 1, point 2 - point orbit while moving SPACECRAFT in orbit, taken as the beginning and the end to calculate the position of the Sun in the field of view of the instrument SETTLEMENT at orientation associated with the SPACECRAFT axes at USC;

αOrb- the angle between the direction of the Sun and the plane of the orbit.

Figure 3, illustrating the position of the orbit and associated SPACECRAFT axes relative to the Sun by the EC is aerialway the Ground plane, given the following notation:

X, Y, Z - linked axis KA;

Point 1, point 2 - point orbit while moving SPACECRAFT in orbit, taken as the beginning and the end to calculate the position of the Sun in the field of view of the instrument SETTLEMENT at orientation associated with the SPACECRAFT axes at USC;

βOrb- the angle between the direction of the Sun and related SPACECRAFT axis X.

The claimed method is implemented as follows.

Up in orbit or in the process of functioning in orbit on-Board digital computing machine (computer) to introduce the orbital parameters at the start of the flight task or by radio from the ground of the complex trajectory measurements.

Calculate the position of the Sun in the field of view of SETTLEMENT for each point of the orbit for the case of orientation associated with the SPACECRAFT axes X, Y, Z, USC (array S1, figure 1).

When the AC output from the shadows (if at the start of the orientation of the SPACECRAFT was in the shadow) set space search apparatus angular velocity (deploy KA) to ensure capture of the Sun field of view of the VILLAGE.

After the capture of the Sun by the POS device and the emergence of information about the current angles α, β the position of the Sun in the field of view of the POS device generated by device PIC, lower (down to 0) search engine angular velocity, ensuring the presence of the Sun in the field of view of the VILLAGE, next produce a reversal of the AC so that the Sun is in the field of view of the app is RA the VILLAGE moved to the desired initial calculated point from the array of S 1(1) the provisions of the Sun in the field of view of the PIC and then continuously deploy the SPACECRAFT from the starting point calculated in the following point.

The coincidence of current angles α, βproduced by the POS device, with the calculated angles α, β n-consistent calculated points (initial plus 2-3 points) associated with the SPACECRAFT axes X, Y, Z are oriented in the direction of the respective axes USC and building a three-axis orientation of the SPACECRAFT in the USC ends, and control is passed to the circuit maintain the orientation of the SPACECRAFT at USC.

The way triaxial orientation of the spacecraft in the orbital coordinate system according to the device orientation to the Sun, characterized in that the onboard computer enter orbit parameters, calculate the Sun position in the field of view of the instrument orientation for each point of the orbit to the orientation associated with the spacecraft axis in the orbital coordinate system, define a space apparatus search engine angular velocity to ensure capture of the Sun field of view of the instrument orientation, then reduce down to zero angular velocity, ensuring the presence of the Sun in the field of view of the instrument orientation, then produced a roll of the spacecraft so that the Sun is in the field of view of device orientation moved to the desired initial designed the th point from the positions of the Sun in the field of view of the device orientation and then continuously deploy the spacecraft from the initial point in the next calculated position of the Sun in the field of view of the instrument orientation for each the point of the orbit at the desired orientation associated with spacecraft axis in the orbital coordinate system.



 

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FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

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9 dwg, 1 tbl

FIELD: cosmonautics, applicable in space activity - space exploration, exploration of the solar system, observation of the Earth from the space, at which it is necessary to determine the space co-ordinates of the space vehicles and the components of their flight velocity vectors.

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2 dwg

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5 dwg, 1 tbl

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3 cl, 13 dwg

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4 cl

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44 dwg

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SUBSTANCE: proposed method includes turning the solar battery panels to working position corresponding to matching of normal to illuminated surface of solar batteries with plane formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also determination of moments of the beginning of solar activity and arrival of high-energy particles onto the spacecraft surface. Then, density of fluxes of said particles is measured and the results are compared with threshold magnitudes. When threshold magnitudes are exceeded, solar battery panels are turned through angle between the said normal and direction to the Sun which corresponds to minimum area of action of particle fluxes on solar battery surfaces at simultaneous supply of spacecraft with electric power. When action of particles is discontinued, solar battery panels are returned to working position. Angle between direction to the Sun and axis of rotation of solar battery panels is measured additionally. In case threshold magnitudes are exceeded, solar battery panels are turned to magnitude of angle between normal to their illuminated surface and direction to the Sun which corresponds to minimum area of action of said particle fluxes on spacecraft surfaces (provided the spacecraft is supplied with electric power). System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System is additionally provided with unit for measurement of angle between direction to the Sun and direction of axis of rotation of solar battery panels, as well as unit for determination of maximum current.

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3 cl, 1 dwg

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3 cl, 1 dwg

FIELD: electric power supply for spacecraft with the aid of solar batteries.

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EFFECT: reduction of negative action of high-energy particle fluxes on solar battery working surface on shaded surface of orbit.

3 cl, 1 dwg

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