Method for orientation of artificial earth satellite

FIELD: physics, control.

SUBSTANCE: invention relates to controlling orientation of artificial earth satellites with solar panels. An artificial earth satellite (3) further includes a self-contained circuit for controlling orientation of the artificial earth satellite relative to the direction towards the sun (2). Upon violation of accuracy of said orientation, orientation of the artificial earth satellite is stopped using an on-board computer simultaneously relative to the direction towards the sun and the earth (1). Said self-contained circuit is turned on, and the solar panels (5) are installed in a fixed position relative to the body of the artificial earth satellite to achieve maximum illumination thereof. Resumption of orientation of the artificial earth satellite using the on-board computer is carried out at a radio command from the earth. The accuracy of orientation of the artificial earth satellite towards the sun can be evaluated from current parameters of the power supply system of the artificial earth satellite. A sign of violation of said orientation can be the beginning of operation of the power supply system in discharge mode of on-board batteries when flying outside shadow portions of the orbit (4).

EFFECT: providing longevity of an artificial earth satellite during prolonged self-contained operation thereof in space.

3 cl, 1 dwg

 

The invention relates to the field of construction of artificial Earth satellites (AES), stabilized by the position of their geometric axes relative to a given coordinate system.

The known method the orientation of the artificial satellite of the Earth, including the orientation of the apparatus relative to the direction of the Sun, the determination of angles between the direction of the Sun and orbital axes coordinate system, the rotation of the device around the direction of the Sun to match the axis of sight of the sensor at the local level with the local vertical of the planet, characterized in that before the rotation direction of the Sun is fixed axis of the apparatus, the projection of the unit vector directions on the axis associated with the device coordinate system is equal to the projections of the unit direction vector to the Sun on the axis of the orbital coordinate system, and turn the apparatus in the plane of the fixing axis - direction to the Sun to match the specified fixed axis of the apparatus with the direction to the Sun (patent No. 2021173, EN).

The closest technical solution is the way the orientation of the artificial Earth satellite, implemented by a satellite with three-axis stabilization of the angular position, comprising a housing having a pitch axis, oriented strictly perpendicular to the orbit plane, and containing at least one surface is, designed mainly for the fact that it was affected by solar radiation, and extending from the housing in a given direction, on-Board computer and connected to a system for determining angular position, designed to determine the angular position of the housing at least relative to the axis of the pitch, active controls the angular position intended for the application of torques, correcting the angular position of the at least pitch, controls, orbit, intended for application to the satellite pulse jet thrust, controls the swing of the specified surface transversely with respect to solar radiation with amplitude, adjustable on-Board computer or from the Ground, and controls the swing located between the said surface and the body of the satellite, characterized in that the said surface passes from the housing parallel to the axis of the pitch, and controls the swing together with the previously mentioned surface are Executive bodies active controls angular position (patent No. 2114770, EN), selected as a prototype.

However, the solution claimed in the known patents do not solve the issue of providing high survivability of the satellite when it occurs during operation of the satellite any with the OEB in the system orientation, including malfunction of the onboard computer.

The task of the claimed invention is the provision of survivability of the satellite during its long-term Autonomous operation in conditions of outer space.

This task is solved in that at orientation artificial Earth satellite using the on-Board computer associated with the system for determining the angular positions of the apparatus relative to the axis of orientation and active management of angular positions, including the orientation of the artificial satellite of the Earth relative to the direction of the Sun and on the Earth, additionally provide an Autonomous attitude control circuit of an artificial Earth satellite with respect to the direction of the Sun and the disruption to the orientation of the artificial Earth satellite in the Sun, the orientation of the artificial satellite of the Earth relative to the direction of the Sun and the Earth using the on-Board computer temporarily cease, this include in the Autonomous loop attitude control of an artificial Earth satellite with respect to the direction of the Sun, with the appropriate installation of solar panels in a fixed position relative to the housing of the artificial Earth satellite to maximize their light, and the subsequent resumption of the orientation is artificial Earth satellite using the on-Board computer carry out radiocommande from the Earth. The accuracy of orientation of the artificial Earth satellite in the Sun assess the current parameters of the power system taking into account the presence/absence of the orbital shadow areas. In addition, the sign of the orientation of the artificial Earth satellite in the Sun take the power in the discharge side of the batteries, in the absence of orbital shadow.

The present invention provides the following advantages.

During operation of the satellite it is exposed to a variety of space factors, such as local static discharge, which can lead to failure of the onboard computer ("hang"). The process of keeping satellites in space will be terminated. Eventually the satellite will begin to lose its original orientation, including the Sun and the Earth.

To monitor this fact can current telemetry system parameters power supply (taking into account the presence/absence of the orbital shadow areas), for example to reduce current solar panels, reducing currents to charge the battery, turn on bit converters and the emergence of currents of the battery discharge.

However, if the satellite is operated in standalone mode (without constant monitoring from the Ground), the disturbance in the orientation of the satellite may is rivetti to full battery discharge, and the loss of the satellite.

Radical protection from emergency situations on the satellite associated with the loss of orientation to the Sun, might be an automatic failover to a backup version of the orientation of the satellite - transition to the Autonomous loop attitude control of an artificial Earth satellite with respect to the direction of the Sun. This mode may not provide the full functionality of the satellite, but it guarantees a positive energy balance that ensures the survivability of the satellite. The transition to an Autonomous control loop orientation should be based on "hard" logic (without using the on-Board computer). The system power supply form the command to switch on the mode of discharge of the rechargeable battery, which is used to automatically transition to an Autonomous control loop orientation. This command is blocked for the time of passage of the satellite orbital shadow area. Lock-unlock can be implemented manually by radio from the Ground and automatically with the on-Board computer. In the latter case must be provided by software measures to unlock, crashes the computer.

During the next communication session with the satellite, in the case of finding it in offline mode orientation, determined by telemetry parameters and eliminate by radio emerged faulty the tee (for example, reset the trip computer) and give the command to return to the nominal orientation of the satellite.

In Fig.1 shows schematically a view of an artificial satellite 3, stabilized on three axes in a pie, for example the geo orbit 4 around the Earth 1.

This satellite contains three axis orientation:

the X - axis, tangent to the orbit and having the same direction as the vector of linear speed of the satellite 3,

- Y axis perpendicular to the orbital plane 4 and oriented in the direction North-South (S-s) Earth 1,

- the Z axis is perpendicular to the axes X and Y, and focused on the Ground 1.

The satellite also contains solar panels, with two wings 5/1 and 5/2, directed respectively to the North and South along its longitudinal axis, here coinciding with the Y axis and oriented relative to the housing around an axis of rotation, approximately coinciding with the Y-axis, under the action of the two drive motors (not shown) that are managed separately. These drive motors are designed to keep the wings BS in the direction perpendicular to the Sun its rays.

When the disruption to the orientation of the satellite 3 Sun 2 the orientation of the satellite relative to the direction of the Sun 2 and direction on the Earth 1 using the onboard computer (not shown) temporarily cease, while included in the Autonomous loop attitude control of satellites 3 relative to N. the Board in the Sun 2 on the X-axis or z-axis. The direction of the two remaining axes can drift and is determined by the current moments of inertia of the satellite. While solar panels 5/1 and 5/2 are constantly focused on the Sun because of those 2 drive motors, which ensure every day (geostationary orbit) the rotation of the solar panels relative to the housing EAS 360 degrees.

The subsequent resumption of the orientation of the satellite 3 using the on-Board computer carry out radiocommande from Earth 1 through command-measuring wireless link (not shown).

The accuracy of the orientation of the satellite 3 Sun 2 assess current parameters of the system power supply (not shown), taking into account the presence/absence of the orbital shadow, a sign of abnormal orientation of the artificial Earth satellite 1 in the Sun take 2 getting started power satellites 3 in the discharge of onboard battery (not shown), in the absence of orbital shadow.

The combination of the two orientations on the same satellite (primary and backup), taking into account the specific recommendations of the transition from one mode of orientation to the other and back, provides additional positive effect of ensuring survivability of the satellite.

Thus, the inventive method the orientation of the artificial Earth satellite provides live which is an artificial satellite with the long-term Autonomous operation of it in terms of space.

1. The way the orientation of the artificial Earth satellite using the on-Board computer associated with the system for determining the angular positions of the apparatus relative to the axis of orientation and active management of angular positions, including the orientation of the artificial satellite of the Earth relative to the direction of the Sun and the Earth, characterized in that it further include an Autonomous attitude control circuit of an artificial Earth satellite with respect to the direction of the Sun and the disruption to the orientation of the artificial Earth satellite in the Sun, the orientation of the artificial satellite of the Earth relative to the direction of the Sun and the Earth using the on-Board computer temporarily cease, this include in the Autonomous loop attitude control of an artificial Earth satellite with respect to the direction of the Sun, with the appropriate installation of solar panels in a fixed position relative to the housing of the artificial Earth satellite to maximize their light, and the subsequent resumption of the orientation of the artificial Earth satellite using the on-Board computer carry out radiocommande from the Earth.

2. The way the orientation of the artificial Earth satellite on p. 1, characterized in that the accuracy of targeting the promotion of artificial Earth satellite in the Sun assess current parameters of its supply system, taking into account the presence/absence of the orbital shadow.

3. The way the orientation of the artificial Earth satellite on p. 1, characterized in that the sign of the orientation of the artificial Earth satellite in the Sun take the power in the discharge side of the batteries, in the absence of orbital shadow areas.



 

Same patents:

Spacecraft // 2540193

FIELD: aircraft engineering.

SUBSTANCE: spacecraft comprises airframe 1 and panel of solar batteries 6 secured at frame 2 shaped to rod-type frame structure shaped to skewed pyramid. Pyramid base 3 is articulated with spacecraft airframe 1 by brackets 5. Said base in initial position is locked by pyro means 4. Pyramid vertex 7 in working position interacts with latch (pos. B) rigidly secured at spacecraft airframe 1. Frame structure 2 and its fasteners feature higher stiffness. This allows increase in frequency and decrease in amplitude of solar battery panels oscillations caused by programmed turns of spacecraft and other manoeuvres.

EFFECT: higher reliability, longer life.

4 dwg

FIELD: transport.

SUBSTANCE: invention relates to control of orientation of a space, in particular, a transport cargo ship (TCS) with stationary solar battery panels (SB). The method includes the rotation of TCS around a normal to the working surface of SB facing towards the Sun with an angular speed of at least 1.5 deg/sec. Meanwhile within the time interval of at least one round the components of the angular speed of TCS in a structural coordinate system are measured. Using the measured values, directions of the main central axes of inertia of TCS are determined. Among these axes an axis other than the axis of the minimum moment of inertia and making the minimum angle with the normal to the working surface of SB is found. The angle between the direction towards the Sun and the plane of the TCS orbit is determined. If this angle exceeds a certain value depending on the specified minimum angle and also - on the minimum and maximum SB currents, TCS is turned. Meanwhile the named found inertia axis is combined with the direction, perpendicular to the orbit plane and making a sharp angle with the direction towards the Sun. TCS is rotated around this axis towards the direction opposite to orbital rotation. During the rotation the current from SB is measured. At the achievement by the current of the minimum value TCS is again turned until the alignment of the named found axis of inertia of TCS with the named perpendicular direction and again the named rotation of TCS is performed.

EFFECT: ensuring of the necessary power return of SB in the mode of TCS rotation around one of its actual main central axes of inertia at maintaining of the axis of the minimum moment of inertia in the orbit plane.

1 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Angle between direction to the Sun and spaceship orbit plane is defined. Spaceship orbit height is defined to determine the half-angle of the Earth disc visible from spaceship. In case said angle exceeds said half-angle the spaceship gravity orientation is constructed at aligning the axis of its minimum moment of inertia that makes the minimum angle with perpendicular to solar battery working surface, with direction to the Earth centre. Spaceship gravity orientation is maintained by spinning it around the axis of minimum moment of inertia at angular velocity defined from the condition of stability of the given gravity orientation of spaceship.

EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.

1 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Spaceship is turned to alignment of the central axis of inertia making the minimum angle with perpendicular to solar battery working surface with direction to the Sun. Spaceship is spun around spaceship around said axis to measure solar battery current. After current reaches minimum permissible magnitudes spaceship is, again, turned to align said axis of inertia with direction to the Sun. Again, spaceship is spun around said axis.

EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.

1 dwg

FIELD: physics; control.

SUBSTANCE: invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval.

EFFECT: high efficiency of the radiator by creating conditions for natural cooling thereof during eclipse of the solar panel for any altitude of an almost circular orbit of the spacecraft.

5 dwg

FIELD: physics.

SUBSTANCE: invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite.

EFFECT: high accuracy of determining navigation-time data on navigation artificial earth satellites by consumers.

4 cl, 12 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to in-flight control over spacecraft equipped with heat radiator and solar battery. Proposed process comprises spacecraft flight in orbit around the planet with solar battery turn to position corresponding to normal to solar battery working surface directed to the Sun. Spacecraft orbital orientation is constructed whereat solar battery spinning plate is parallel with spacecraft orbit plane while solar battery is located on the Sun side relative to orbit plane. Spacecraft orbit altitude and angle between direction to the Sun and spacecraft orbit plane are defined. Magnitude of said angle (β*) is defined whereat duration of turn shadow section equals the necessary time of radiator heat release in said turn. Orbit turns are defined wherein current magnitude of said angle is larger than β*. In said turns, solar battery is turned around crosswise and lengthwise rotation axes unless shadowing of solar battery radiator. Note here that minimum departure of orientation of solar battery working surface to the Sun. Spacecraft orbital flight is conducted in near-circle orbit at altitude not exceeding a definite design value.

EFFECT: higher efficiency of radiator with solar battery shadowed at whatever position of spacecraft on orbit turn.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to spacecraft (SC) motion control using solar radiation pressure forces distributed over SC working zones. The latter are formed as flat parallel optically transparent droplet flows. Distance between droplets of R radius in each flow in its lengthwise direction (Sx) and frontal-lateral direction (Sy) is divisible by 2R. Number of flows is n=(Sx/2R)1. By mutual bias of flows in direction of their motion for 2R distance droplet mist flows are generated in number of m=(Sy/2R)1. Each of the mentioned flows is biased relative to previous flow for 2R distance in frontal-lateral direction. Thus opacity in frontal-lateral direction and transparency in direction of plane perpendicular to a flow is created. Unit distributed light pressure force is regulated by changing radius and number of droplets coming to point of it application in unit time. Total action value is regulated by changing number of droplet jets.

EFFECT: increase of efficiency of light pressure distributed external forces usage by means of decreasing their disturbing effect on relative SC motion.

3 dwg, 1 tbl

FIELD: electricity.

SUBSTANCE: result is achieved by increasing strength of connection of shunting diodes and solar elements, increased repeatability of the process of manufacturing of the solar battery of spacecrafts due to optimisation of the technology of manufacturing of shunting diodes and solar elements of the solar battery, and also switching buses that connect the solar elements and shunting diodes, which are made as multi-layer. The solar battery for small-size spacecrafts comprises the following: panels with modules with solar elements (SE) adhered to them, a shunting diode; switching buses that connect the face and reverse sides of the shunting diodes with solar elements, at the same time the shunting diode is installed in the cut in the corner of the solar element, at the same time switching buses are made as multi-layer, made of molybdenum foil, at two sides of which there are serial layers of vanadium or titanium, a layer of nickel and a layer of silver, accordingly.

EFFECT: increased resistance of solar batteries to thermal shocks, to impact of mechanical and thermomechanical loads, increased manufacturability of design, extended active life of a solar battery of spacecrafts, increased functional capabilities due to expansion of temperature range of functioning and optimisation of design of a solar battery, simplification of a switching system.

7 cl, 4 dwg, 3 tbl

FIELD: transport.

SUBSTANCE: invention relates to spacecraft electric power supply with the help of solar batteries. Proposed method comprises definition of preset angle of solar battery orientation to the Sun by measured angular position of normal to battery working surface and computation of design angle relative thereto. Solar battery is spinned in direction of decrease in mismatch between preset and design angles. Solar battery acceleration angle (αAC) and deceleration angle (αDEC) are defined. Design angle is corrected when angle transducer readings vary by discrete sector of solar battery turn. Threshold of operation and drop-away (αT) and (αD) are set to terminate battery spinning if mismatch between preset angle and current angle increases but not over αT. Solar battery angular velocity is set or the order and larger than maximum angular velocity of spacecraft revolution around the Earth while discrete sector magnitude is set to smaller than αT. Solar battery working angle (αW) is set provided that αT < αW < (α"ГОР" - 2·(αAC + αDEC)). Angular position of closest beam of angle αW is assigned to preset angle if direction to the Sun in projection to the plane of spinning of said normal is located outside of αW. Is angular position of said normal is outside αW to vary in direction of increase of angle relative to nearest beam of angle αW, failure warning is generated to terminate control over solar battery.

EFFECT: ruled out jamming and breakage of solar battery panels or spacecraft onboard hardware at turns from 90° to 180°.

3 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention can be used for control over program turn of accelerating unit with the help of fixed constant-thrust engines of orientation. Angular velocity is increased at acceleration and inertial flight and decreased to zero at deceleration and pulsed initiation of orientation engines. Level of fuel component in the tank that brings about the most tangible effect on turn dynamics is measured Angle mismatch and acceleration unit angular velocities are intermittently measured at turn as well as deflection of said fuel level from acceleration unit lengthwise axis Orientation engine are shut down at the ends of acceleration path and switched on at deceleration start path.

EFFECT: acceleration unit turn with damping of fuel components oscillations at ramming engine operation path section.

6 dwg

FIELD: transport.

SUBSTANCE: invention relates to control of orientation of a space, in particular, a transport cargo ship (TCS) with stationary solar battery panels (SB). The method includes the rotation of TCS around a normal to the working surface of SB facing towards the Sun with an angular speed of at least 1.5 deg/sec. Meanwhile within the time interval of at least one round the components of the angular speed of TCS in a structural coordinate system are measured. Using the measured values, directions of the main central axes of inertia of TCS are determined. Among these axes an axis other than the axis of the minimum moment of inertia and making the minimum angle with the normal to the working surface of SB is found. The angle between the direction towards the Sun and the plane of the TCS orbit is determined. If this angle exceeds a certain value depending on the specified minimum angle and also - on the minimum and maximum SB currents, TCS is turned. Meanwhile the named found inertia axis is combined with the direction, perpendicular to the orbit plane and making a sharp angle with the direction towards the Sun. TCS is rotated around this axis towards the direction opposite to orbital rotation. During the rotation the current from SB is measured. At the achievement by the current of the minimum value TCS is again turned until the alignment of the named found axis of inertia of TCS with the named perpendicular direction and again the named rotation of TCS is performed.

EFFECT: ensuring of the necessary power return of SB in the mode of TCS rotation around one of its actual main central axes of inertia at maintaining of the axis of the minimum moment of inertia in the orbit plane.

1 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Angle between direction to the Sun and spaceship orbit plane is defined. Spaceship orbit height is defined to determine the half-angle of the Earth disc visible from spaceship. In case said angle exceeds said half-angle the spaceship gravity orientation is constructed at aligning the axis of its minimum moment of inertia that makes the minimum angle with perpendicular to solar battery working surface, with direction to the Earth centre. Spaceship gravity orientation is maintained by spinning it around the axis of minimum moment of inertia at angular velocity defined from the condition of stability of the given gravity orientation of spaceship.

EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.

1 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to aerospace engineering. Proposed method comprises supply spaceship spinning around perpendicular to solar battery working surface directed to the Sun at angular velocity of at least 1.5 degree/s. During said spinning at time interval of duration making at least one circuit supply angular velocity components are measured in structural system of coordinates. Measured magnitudes are used to define the directions of the main central axis of inertia of supply spaceship. Spaceship is turned to alignment of the central axis of inertia making the minimum angle with perpendicular to solar battery working surface with direction to the Sun. Spaceship is spun around spaceship around said axis to measure solar battery current. After current reaches minimum permissible magnitudes spaceship is, again, turned to align said axis of inertia with direction to the Sun. Again, spaceship is spun around said axis.

EFFECT: higher power yield of solar batteries owing to radiation reflected from the Earth at spaceship gravity orientation with spinning with allowance for actual main central axes of inertia.

1 dwg

FIELD: transport.

SUBSTANCE: invention relates to aerospace engineering and can be used in spacecraft engines. Power plant comprises cryogenic tank with shield-vacuum heat insulation and channel with heat exchanger, flow control valve, booster pump, intake with capillary accumulator with heat exchanger and throttle and hydropneumatic system with pipeline. Channel cross-section sizes comply with maximum outer sizes of heat exchanger cross-section.

EFFECT: cooling of cryogenic component in capillary accumulator.

3 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Spacecraft velocity in atmosphere at initial flight part increases (spacecraft flies toward conditional orbit pericentre). Atmosphere density is low yet to cause notable spacecraft deceleration. As spacecraft reaches atmosphere dense layers its velocity decreases to reach atmosphere enter velocity for angle of roll (γ) γ=π to be changed to γ=0. This manoeuvre allows changing the spacecraft to flight part with maximum aerodynamic performances. In flight with γ=0 continuous skip path is maintained whereat spacecraft velocity decreases monotonously. Maximum skip height reached, angle of attack o spacecraft increases, hence, spacecraft intensive deceleration occurs.

EFFECT: decreased final velocity at soft landing system operation, fuel savings.

1 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Proposed method consists in selection of conditions for changing the angle of roll to zero at changing the spacecraft from isothermal descent section (IDS) to skip path. With spacecraft in IDS, angle of roll (γ) is, first, increased to decrease aerodynamic performances and to maintain constant temperature at critical area of spacecraft surface. As flight velocity decreases angle (γ) is decreased from its maximum. In IDS, increase in aerodynamics does not cause further temperature increase over its first peak. Therefore selection of the moment of changing to γ=0 allows efficient deceleration of spacecraft at the next step of flight. The best option is the descent of spacecraft of IDS when γ reaches its maximum. Here, angle of attack is set to correspond to maximum aerodynamic performances. This increases the duration of final flight stage and deceleration efficiency. Increase in angle of attach after descent from IDS and completion of climb results in increased in drag, hence, decrease in velocity at initiation of soft landing system.

EFFECT: minimised final velocity and maximum temperature at surface critical area, lower power consumption.

2 dwg

FIELD: physics; control.

SUBSTANCE: invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval.

EFFECT: high efficiency of the radiator by creating conditions for natural cooling thereof during eclipse of the solar panel for any altitude of an almost circular orbit of the spacecraft.

5 dwg

FIELD: transport.

SUBSTANCE: invention relates to space cable systems (SCS) and can be used for the transfer of SCS to a spinning mode in the orbit plane without the application of jet engines. SCS development is executed from its initial compact state in the circular orbit by the repulsion of objects at a low relative speed. SCS end weights are connected by a cable, its length being varied by a cable feed-haul-in device arranged on one of the end objects. The objects are separated by a vector of local peripheral speed, for example, by a pusher. The objects are driven by a start pulse to separate the objects in practically free paths at the free feed of the cable. The cable development is terminated by the SCS transfer to a stable mode of associated pendulum motion at the stretched preset-length cable. At a definite range of angular phases of this mode the SCS objects are stretched by hauling in the cable at a definite constant speed. This results in changing the SCS into the spinning mode at a preset power integral and fixed final end of the cable.

EFFECT: relaxed weight-size constrictions of SCS, enhanced performances.

8 dwg

FIELD: physics, navigation.

SUBSTANCE: group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections.

EFFECT: improvement of orientation accuracy and operational reliability in case of failures of orientation angle sensor and sensor of angular velocity of space vehicle rotation.

2 cl, 2 dwg

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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