Method of three-axis orientation of spacecraft in orbital coordinate system |
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IPC classes for russian patent Method of three-axis orientation of spacecraft in orbital coordinate system (RU 2247684):
     
 Single-stage axial-symmetric reusable booster and a way to create a lifting force during the movement of single-stage booster / 2226169
The invention relates to rocket and space technology and can be used to create launch vehicles (LV), including conversion, for a spacecraft in low earth orbit
 The method of forming the terminal management guidance upper stage into orbit / 2223894
The invention relates to space technology, and more particularly to management of orbital maneuvers boosters with lively marching rocket engines
 The control system turns spacecraft / 2213681
The invention relates to automatic control systems nonstationary, mainly space objects
 The control system turns spacecraft / 2213029
The invention relates to automatic control systems nonstationary, mainly cosmic objects
 The control system turns spacecraft / 2211788
The invention relates to automatic control systems nonstationary, mainly cosmic objects
 The control system turns spacecraft / 2211787
The invention relates to automatic control systems nonstationary, mainly space objects
 The correction method, the parameters of change of longitudinal movement in terminal management guidance upper stage into orbit / 2211786
The invention relates to space technology, and more particularly to the onboard controls boosters with lively marching rocket engines
 The way to determine when the end time of the maneuver and the sustainer engine cutoff based on the numerical prediction of the motion of the upper stage, / 2209159
The invention relates to space technology, and more particularly to an onboard means of terminal control boosters with unregulated lively marching rocket engines
 The method of identifying conditional combustion time of the mass of the upper stage / 2209158
The invention relates to space technology, and more particularly to an onboard means of terminal control boosters with unregulated lively marching rocket engines
 The identification system conventional combustion time of the mass of the upper stage / 2209157
The invention relates to space technology, and more particularly, to an onboard means of terminal control boosters with unregulated lively marching rocket engines
 Method of three-axis orientation of spacecraft in orbital coordinate system / 2247684
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.
 Method of correction of parameters of longitudinal motion change program at terminal control of cryogenic stage guidance on preset orbit / 2254271
Parameters of motion of cryogenic stage at moment of cruise engine cutoff are predicted and radius of deviation of radius and radial velocity of cryogenic stage from their preset magnitudes are determined. Signals for correction of pitch angle and rate of pitch are shaped for compensation of said deviation. Pitch angle correction signal is limited at preset level and its excess above this level is determined. When signal is shaped for limitation, addition to correction pitch rate signal is formed. This correction is equal to product of said excess by ratio of functions of sensitivity of radial velocity of cryogenic stage to pitch angle and rate of pitch. Resultant pitch rate correction signal is formed as sum of this signal determined without taking into account pitch angle correction signal limitation and addition. Thus, priority follow-up of velocity error is ensured at limited pitch angle correction.
 Method for location of space vehicles / 2275650
The method consists in the fact that in the intermediate orbit simultaneously with determination of the co-ordinates of the space vehicle (SV)
 Method of forming program for orientation of cryogenic stage at terminal control of injection into preset orbit / 2282568
Swivel combustion chamber of cruise engine is used for angular orientation and stabilization of cryogenic stage of spacecraft. Proposed method includes predicting parameters of motion of cryogenic stage at moment of cut-off of cruise engine; deviation of radius and radial velocity from preset magnitudes are determined; angle of pitch and rate of pitch are corrected and program of orientation of thrust vector for subsequent interval of terminal control is determined. By projections of measured phantom accelerations, angle of actual orientation of cruise engine thrust vector and misalignment between actual and programmed thrust orientation angles are determined. This misalignment is subjected to non-linear filtration, non-linear conversion and integration. Program of orientation of cryogenic stage is determined as difference between programmed thrust orientation angle and signal received after integration. Proposed method provides for compensation for action of deviation of cruise engine thrust vector relative to longitudinal axis of cryogenic stage on motion trajectory.
 Method of control of cluster of satellites in geostationary orbit (versions) / 2284950
Proposed method consists in measurement of parameters of satellite orbits, determination of orbital elements, comparison of them with required ones and performing of correcting maneuvers with the aid of thrusters. Satellite inclination vectors are brought to circular areas of their permissible change which are spaced apart so that angle between line connecting the end of vector with center of its circular area and direction to Sun should exceed right ascension of Sun by 180°. According to first version, vectors of satellite eccentricity are shifted to similar circular areas so that similar line lags behind direction to the Sun by half angular displacement of vector over circumference of its natural drift within circular area. Then, distances between satellites are changed within required limits compensating for quasi-secular increment of inclination vector and correcting eccentricity vector so that at passing the center of interval between point of circumference entry of its natural drift to its circular area and point of exit from this area, line connecting the center of this circumference and center of circular area coincide with direction to the Sun. In case circular area of permissible change of each eccentricity vector is close to circumference of its natural drift (second version), said line for this area is matched with direction to the Sun and no correction is made in this case.
 Method for missile take-off from aircraft for orbit injection of payload / 2289084
The method consists in separation of the missile with a payload from the carrier aeroplane and its transition to the state with initial angular parameters of motion in the vertical plane. After separation the missile is turned with the aid of its cruise engine, preliminarily using the parachute system for missile stabilization. The parachute system makes it possible to reduce the duration of the launching leg and the losses in the motion parameters (and the energy) in this leg. To reduce the missile angular bank declination, the strand of the parachute system fastened in the area of the missile nose cone is rehooked. To reduce the time of missile turning towards the vertical before the launcher, the cruise engine controls are preliminarily deflected to the preset angles and rigidly fixed. By the beginning of missile control in the trajectory of injection this fixation is removed. In the other modification the missile turning is accomplished by an additional jet engine installation. It is started depending on the current angular parameters of missile motion so that by the beginning of controlled motion in the trajectory of injection the missile would have the preset initial angular parameters of motion.
 Self-contained onboard control system of "gasad-2a" spacecraft / 2304549
Proposed system includes control computer, star sensor, Earth sensor, storage and timing device, processors for control of attitude, processing angular and orbital data, inertial flywheels and spacecraft orbit correction engine plant. Used as astro-orienters are reference and navigational stars from celestial pole zone. Direction of spacecraft to reference star and direction of central axis of Earth sensor to Earth center are matched with plane formed by central axes of sensors with the aid of onboard units. Shift of direction to reference star relative to central axis of Earth sensor is considered to be latitude change in orbital position of spacecraft. Turn of navigational star around reference star read off sensor base is considered to be inertial longitude change. Point of reading of longitude is point of spring equinox point whose hour angle is synchronized with the board time. This time is zeroed upon completion of Earth revolution. Stochastic measurements by means of static processing are smoothed-out and are converted into geographic latitude and longitude parameters. Smoothed inertial parameters are compared with parameters of preset turn of spacecraft orbit found in storage. Revealed deviations of orbit are eliminated by means of correction engine plant.
 Method of control of spacecraft solar battery position and system for realization of this method / 2322372
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.
 Method of control of spacecraft solar battery position and system for realization of this method / 2322373
Proposed method includes turning of solar batteries to the working position corresponding to matching of normal to their illuminated surface with plane formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also measurement of density of fluxes of solar electromagnetic radiation and high-energy particles determining the moments of beginning of solar activity and arrival of said particles to spacecraft surface. Additional measurement includes determination of appearance of signs of negative action of particle flux on spacecraft. During these moments, onboard solar batteries are charged to maximum level. When density of particle flux exceeds threshold magnitude, solar battery panels are turned through angle between said normal and direction to the Sun corresponding to minimum action of particle fluxes on solar battery surfaces. Discharge of storage batteries is hoped to close the energy gap on board the spacecraft. At minimum permissible level of storage battery charge, storage batteries are disconnected from load. When action of particles on spacecraft is discontinued, solar battery panels are returned to working position. System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System is provided with unit for determination of current from solar batteries, unit for determination of moments of appearance of signs of negative action of high-energy particles on spacecraft and unit for setting the permissible level of charge of storage batteries.
 Method of control of spacecraft solar battery position and system for realization of this method / 2322374
Proposed method includes turning the solar battery panels to working position corresponding to matching of normal to their illuminated surface formed by axis of rotation of solar battery panels and direction to the Sun. Proposed method includes also measurement of density of fluxes of solar electromagnetic radiation and high-energy particles followed by determination of moments of beginning of solar activity and arrival of high-energy particles to spacecraft surface. Method includes additionally measurement of spacecraft orbit altitude and angle between direction to the Sun and plane of spacecraft orbit. In case density of particle flux exceeds threshold magnitudes, solar battery panels are turned on illuminated surface of spacecraft orbit through angle (αs min) between said normal and direction to the Sun corresponding to minimum area of action of particle fluxes on spacecraft surfaces at supply of spacecraft with required amount of electric power. On shaded side of orbit, solar batteries are turned from direction of particle flux through maximum angle. When spacecraft escapes from shadow, reverse turn of solar battery panels is completed through said angle αs min. Upon completion of action of particle flux on spacecraft, solar battery panels are returned to working position. System proposed for realization of this method includes units and their couplings for performing the above-mentioned operations. System includes additionally unit for determination of intensity of spacecraft illumination, unit for measurement of spacecraft orbit altitude, unit for measurement of angle between direction to the Sun and spacecraft orbital plane, unit for control of turn of solar battery to position opposite to direction to the Sun, NO-gate and switch.
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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
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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